About Transonic Compressor Tandem Design: A Principle Study

The development of modern axial compressors has already reached a high level. Therefore an enlargement of the design space by means of new or advanced aerodynamic methods is necessary in order to achieve further enhancements of performance and efficiency. The tandem arrangement of profiles in a transonic compressor blade row is such a method. For an efficient industrial application the knowledge of the fundamental design principles is needed. This paper presents the recent research work on transonic compressor tandem profiles at DLR Institute of Propulsion Technology. It deals with the fundamental description of the operation principles of a modern transonic compressor tandem cascade. By considering these principles and based on an optimization database with over 1200 members design recommendations are developed.

Numerical Analysis on Axial Compressor Stage Performance With Vortex Generators

The performance of axial flow compressor stage can be improved by minimizing the effects of secondary flow and by avoiding flow separation. At higher blade loading, interaction of tip secondary flow and separated flow on blade surface is more near the tip of the rotor. This results in stall and hence decreases compressor performance. A previous study [1] was carried out to improve the performance of a rotating row of blades with the help of Vortex Generators (VGs) and considerable effects were observed. The current investigation is carried out to find out the effect of Vortex Generator (VG) on the performance of a compressor stage. NASA Rotor 37 with NASA Stator 37 (stage) is used as a test case for the current numerical investigation. VGs are placed at different chord wise as well as span wise locations. A mesh sensitivity study has been done so that simulation result is mesh independent. The results of numerical simulation with different geometrical forms and locations of VGs are presented in this paper. The investigation includes a description of the secondary flow effect and separation zone in compressor stage based on numerical as well as experimental results of NASA Rotor 37 with Stator 37 (without VG). It is also observed that the shape and location of the VG impacts the end wall cross flow and flow deflection [1], which result in enhanced stall range.

Closed-Loop Active Flow Control of the Wake of a Compressor Blade by Trailing-Edge Blowing

This paper presents a closed-loop active flow control strategy to reduce the velocity deficit of the wake of a compressor stator blade. The unsteady stator-rotor interaction, caused by the incoming stator wakes, generates fast changes of the rotor blade loading, affecting the stability and the performance of the overall compressor. Negative effects will be seen likewise when unsteady combustion concepts, such as a pulsed detonation, produce upstream disturbances. Furthermore, the periodic unsteady flow leads to additional undesired effects such as noise and blade vibrations. A controlled reliable manipulation of the stator wake is a way to handle these issues. Therefore, investigations on wake manipulation with trailing-edge blowing were carried out on a new low-speed cascade test rig. Detailed information about the wake profile is obtained by five-hole probe measurements in a plane downstream of the cascade for the natural and the actuated flow at a Reynolds number of 6×105. These measurements show a significant reduction of the wake velocity deficit for the investigated actuator geometry with an injection mass flow of less than 1% of the passage mass flow. Based on these results a position in the wake was chosen which is representative for the actuation impact on the velocity deficit. There, a hot-wire-probe measurement serves as the controlled variable. A family of linear dynamic black-box models was identified from experimental data to account for nonlinear and unmodelled effects. Static nonlinearitiy was compensated for by a Hammerstein model to reduce the model uncertainty and get a higher controller performance. To handle off-design conditions, a robust controller working in a range of Reynolds numbers from 5×105 to 7×105 was synthesized. The task of the controller is to rapidly regulate the controlled variable to a reference velocity by changing the blowing amplitude. The synthesized robust controller was successfully tested in closed-loop experiments with good results in reference tracking for pulse series up to 20 Hz. This translates into a much higher frequency when scaled to the dimension of a real machine.

Influence of Turbulent Flow Characteristics and Coherent Vortices on the Pressure Recovery of Annular Diffusers: Part A — Experimental Results

Exhaust diffusers downstream of turbines are used to transform the kinetic energy of the flow into static pressure. The static pressure at the turbine outlet is thus decreased by the diffuser, which in turn increases the technical work as well as the efficiency of the turbine significantly. Consequently, diffuser designs aim to achieve high pressure recovery at a wide range of operating points. Current diffuser design is based on conservative design charts, developed for laminar, uniform, axial flow. However, several previous investigations have shown that the aerodynamic loading and the pressure recovery of diffusers can be increased significantly if the turbine outflow is taken into consideration. Although it is known that the turbine outflow can reduce boundary layer separations in the diffuser, less information is available regarding the physical mechanisms that are responsible for the stabilization of the diffuser flow. An analysis using the Lumley invariance charts shows that high pressure recovery is only achieved for those operating points in which the near-shroud turbulence structure is axi-symmetric with a major radial turbulent transport component. This turbulent transport originates mainly from the wake and the tip vortices of the upstream rotor. These structures energize the boundary layer and thus suppress separation. A logarithmic function is shown that correlates empirically the pressure recovery vs. the relevant Reynolds stresses. The present results suggest that an improved prediction of diffuser performance requires modeling approaches that account for the anisotropy of turbulence.

Numerical Investigation of a Cantilevered Compressor Stator at Varying Clearance Sizes

The clearance size of cantilevered stators affects the performance and stability of axial compressors significantly. Numerical calculations were carried out using the commercial software FINE/Turbo for a 2.5-stage highly loaded transonic axial compressor, which is of cantilevered stator for the first stage, at varying hub clearance sizes. The aim of this work is to improve understanding of the impact mechanism of hub clearance on the performance and the flow field in high flow turning conditions. The performance of the front stage and the compressor with different hub clearance sizes of the first stator has been analyzed firstly. Results show that the efficiency decreases as clearance size varies from 0 to 3% of hub chordlength, but the operating range has been extended. For the first stage, the efficiency decreases about 0.5% and the stall margin is extended. The following analysis of detailed flow field in the first stator shows that the clearance leakage flow and elimination of hub corner separation is responsible for the increasing loss and stall margin extending respectively. The effects of hub clearance on the downstream rotor have been discussed lastly. It indicates that the loss of the rotor increases and the flow deteriorates due to increasing of clearance size and hence the leakage mass flow rate, which mainly results from the interaction of upstream leakage flow with the passage flow near pressure surface. The affected region of rotor passage flow field expands in spanwise and streamwise direction as clearance size grows. The hub clearance leakage flow moves upward in span as it flows toward downstream.

Iterative Learning Active Flow Control Applied to a Compressor Stator Cascade With Periodic Disturbances

This contribution presents the capability of iterative learning active flow control to decrease the impact of periodic disturbances in an experimental compressor stator cascade with sidewall actuation. The periodic disturbances of the individual passage flows are generated by a damper flap device that is located downstream of the trailing edges of the blades. These mimic the throttling effect of periodically closed combustion tubes in a pulsed detonation engine. For the purpose of rejecting this disturbance the passage flow is manipulated by fluidic actuators that introduce an adjustable amount of pressurized air through slots in the sidewalls of the cascade. Pressure sensors that are mounted flush to the suction surface of the middle blade provide information on the current flow situation. This data is fed back in real-time to an optimal iterative learning controller. By learning from period to period the controller modifies the actuation amplitude such that, eventually, a control command trajectory is calculated that reduces the impact of the periodic disturbance on the flow in an optimal manner.

Development and Aerothermal Investigation of Integrated Combustor Vane Concept

This paper presents the development and aerothermal investigation of the Integrated Combustor Vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010 [1], the conventional Nozzle Guide Vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept is developed using the inhouse CFD code TBLOCK. Aerothermal experiments are conducted using a modular high speed linear cascade, designed to model the flow at the combustor-vane interface. The facility comprises two can combustor transition ducts and either four Conventional Vanes (CVs) or two Integrated Vanes (IVs). The experimental study validates the linear CFD-simulations of the IV development. Annular full stage CFD-simulations, used to evaluate aerodynamics, heat transfer and stage efficiency, confirm the trends of the linear numerical and experimental results and thus demonstrate the concept’s potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile- and endwall-loss. Combined with an improved rotor performance these aerodynamic benefits result in a gain in stage efficiency of above 1%, illustrated by unsteady stage simulations. A distinct reduction in HTC levels on vane surfaces, in the order of 25%–50%, and endwalls is observed and attributed to an altered state of boundary layer thickness. The development of IV’s endwall- and LE-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

Desensitization of Axial Compressor Performance and Stability to Tip Clearance Size

This paper presents a computational and analytical study to identify and elucidate fundamental flow features associated with the desensitization of performance and aerodynamic stability of an axial compressor rotor to tip clearance change. Parametric studies of various design change on a baseline double circular arc axial rotor led to the identification of two flow features associated with reducing sensitivity to tip clearance, namely high incoming meridional momentum in the tip region and reduction/elimination of double tip leakage. Numerical experiments were subsequently performed on the baseline rotor geometry to validate these two flow features and explain the associated flow physics by variations in incoming meridional momentum and pitch size. Finally, two designs were proposed, namely full forward chordwise sweep and partially low stagger angle, to exploit these flow features. The results indicated that both designs produce the intended flow effects and exhibit lower performance and aerodynamic stability sensitivity to tip clearance.

Aerodynamic Optimization of the Gas Outlet Casing of the Axial Turbine Stage in a Turbocharger With Experimental Validation

This paper presents a design process for optimizing the aerodynamic performance of the gas outlet casing on single-stage turbines with high specific flow capacity. A full-annulus flange-to-flange (inlet-to-outlet) steady-state CFD model of the turbine stage which takes into account the interaction between the rotor and the gas outlet casing and ensures an accurate inflow condition for the latter, is used to predict the turbine stage performance. A meta-model based optimization for the gas outlet casing is then performed and a simplified CFD model is used for sampling and optimization. The geometry of the gas outlet casing is fully parameterized to enable the simultaneous variation of diffuser and collector geometry and an ordinal regression optimization algorithm is adopted for the objectives of maximizing the static pressure recovery of the gas outlet casing and ensuring the design robustness. Extensive test measurement of the turbine stage with its baseline and optimized gas outlet casing geometries on a full scale turbocharger test bench validates the CFD results and confirms the significant improvement of the exhaust casing pressure recovery, which leads to an improvement of turbine efficiency between 1.3 and 2.4 % points over the relevant considered operating range. Traverse measurement using five-hole probes and the flow field predicted by CFD are in good agreement. Evaluation of the CFD results highlights a significant loss in the collector despite a high pressure recovery at the end of the diffuser for the baseline gas outlet casing. For the optimized geometry, the more uniform flow at the diffuser outlet results in greatly reduced loss in the 90° turn in the casing, and thus higher pressure recovery and turbine efficiency.

Some Experiences About the Impact of Unsteadiness in Turbine Flows

This paper describes some experiences about impact of unsteadiness in turbine flows, with a special focus on the effects of potential interaction on aerodynamic performance. The main motivation consists in trying to identify some design areas in which some further margins of improvement could be found, provided the designer chooses the proper computational framework. The underlying idea is that the approximations associated with the steady-state picture of a turbine stage might prevent the designer from unlocking the full potential of the stage, especially when the design requirements imply a challenging aerodynamics. To this end, three common design topics are presented in which the step from the classical steady-state approach to the time-accurate one unveils relevant issues, which in turn have an impact on aerodynamic performance: stator/rotor interaction in transonic stages, the choice of the axial gap between stator and rotor, and the choice of the blade count ratio. In all reported cases, significant departures are found between steady and time-averaged results, and the basic fluid mechanisms responsible for them are examined. In particular, an attempt is made to emphasize limitations deriving from of the steady-state picture of the turbine flow field, in order to warn the designer about the possible traps of the steady-state assumption.