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.

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.

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.

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.

Requirements and Limitations of Controlling Turbine Tip Shroud Cavity Flow Mixing With Bladelets on Rotating Shroud

A potential means of significantly reducing the cavity exit mixing loss, a dominant primary loss mechanism in turbine tip shroud cavity flow, is assessed. The operational constraints on the turbine stage dictate that losses may only be mitigated through configuration changes within the cavity. A configuration, known herein as the Hybrid Blade, features a shrouded main blade with a row of high aspect ratio bladelets affixed to the rotating shroud is formulated and shown to nearly eliminate the cavity exit mixing loss. However the Hybrid Blade configuration incurs a penalty associated with bladelet low Reynolds number effects, cavity inlet flow asymmetry introduced by the scalloped shroud, and a resulting mismatch with the upstream vane as well as downstream diffuser. This penalty offsets the efficiency gain from mitigating cavity exit mixing loss. For the Hybrid Blade system, it can thus be inferred that the turbine stage and the diffuser need to be reconfigured to accommodate the modified tip shroud, and the bladelets redesigned for low Reynolds number operation and cavity inlet flow asymmetry to achieve an overall benefit.

Design of an Air-Starter Turbine and Starting Performance Prediction Through the Numerical Analysis

To get the required power, a procedure for starting an engine is required. For an engine of the large ship or land-based power plant, two kinds of mechanisms exist: one is to use an electric motor and the other air-starter. As one of the starting methods of heavy-duty engines, the air-starter gives rotating power from the blade lifting force by the compressed air flow through the turbine to rotate the stopped engine. Since the air turbine has a large weight-to-power ratio, simple structure and is spark-free, many heavy-duty engine are equipped with this device for their starting process. In order to get high enough power required to rotate the engine up to sufficient speed in a short time, high instantaneous torque is required. An impulsive supersonic turbine can be applied for this purpose. In this case, an one-stage air turbine with a minimum length convergent-divergent nozzle and transition arc shaped rotor was utilized to reduce shock loss. The performance of the desired turbine is evaluated numerically and compared with experimental results. The spin-down test is performed to estimate the performance of the designed turbine and numerical analysis is performed using the commercial tool ANSYS CFX.

Aerodynamic Implications of Reduced Vane Count

Given the shortage of fossil fuels and the growing greenhouse effect, one strive in modern gas turbines is to make maximum usage of the burnt fuel. By reducing the number of vanes or blades and thereby increasing the loading per vane (or blade) it is possible to spend less cooling air, which will have a positive impact on the combined cycle efficiency. It also reduces the number of components and usage of metal and thereby also the cost of the engine. These savings should be achieved without any efficiency deficit in aerodynamic efficiency. Based on the fact, aerodynamic investigations were performed to see the aerodynamic implications of reduced vane number in a transonic annular sector cascade. The number of new nozzle guide vane was reduced with 24% compared to a previous design with higher vane count. The investigated vanes were two typical high pressure gas turbine vanes. Results regarding the loading indicated an expected increase with the reduced vane case. The minimum static pressure at the suction side is lower and at an earlier location for the reduced vane case and therefore, an extension of the trailing edge deceleration zone is observed for the reduced vane case. Results regarding losses indicate that even though the losses produced per vane significantly increases for the reduced vane case, a comparison of mass averaged losses between the reduced vane case and previous vane case show similar spanwise loss distributions. Assessing results leads to a conclusion that the reduction of the number of vanes in the first stage seems to be a useful method to save cooling flow as well as material costs without any significant deficit in overall efficiency.