Rotating stall control for axial flow compressors

Non-axisymmetric end wall profiling is now a well established design methodology in axial flow turbines, used principally to improve their aerodynamic efficiency by reducing secondary loss. However, profiled end walls (PEWs) have yet to find an in-service application in a gas turbine compressor. This two-part paper presents the results of a number of studies, both experimental and computational, into the potential aerodynamic benefits of applying PEWs in axial flow compressors. The second paper describes an investigation into PEWs as a means of suppressing stator hub corner stall. An in-house design optimisation system (SOPHY) was applied to a typical embedded stator row in a multi-stage HP compressor. This enabled a new PEW geometry to be defined which is shown to be an effective alternative to 3-D aerofoil shaping in controlling stator hub corner stall. The results of this new study and of the work described in the first paper have been combined in a further investigation. PEWs were applied to an HPC with purely 2-D blading, which exhibited extensive stator hub corner stall at off-design conditions. CFD analysis showed that stator hub corner stall was successfully suppressed by the PEWs. It is concluded that PEWs can significantly affect the end wall flow field and have similar effects on corner stall as 3-D blading. The potential for improving the performance of axial flow compressors using PEWs is discussed. To successfully achieve any such improvements, extensive computational resources that are dedicated to design optimisation will be required.

This paper investigates the instability inception phenomenon in a variable rotor tandem (VRT) stage compressor by developing a numerical model based on unsteady Reynolds-averaged Navier–Stokes (URANS) equations. The main objective is to assess the idea of replacing a combination of one inlet guide vane row and one conventional axial-flow stage with a VRT stage to reduce the weight and size of axial-flow compressors. Steady performance analysis of a tandem stage axial-flow compressor, designed and constructed in the Dana laboratory of the Amirkabir University of Technology, was previously reported. In that study, it was discovered that at both design and off-design speeds, a VRT stage is more effective than a variable stator tandem stage in postponing the surge point and extending the safe operation range. However, Reynolds-averaged Navier–Stokes (RANS) modeling did not show the origin of instability inception and particulars of such highly complex flow phenomena. To explore the physics of flow more accurately, a full-annulus three-dimensional URANS model is developed to study near-stall performance. Transient simulations at mass flow rates less than the stall point show that instability originates from hub corner stall formation on the suction surface of the front stator tandem blade at about 0.97 revolution time. In 4.62 revolutions, the stall zone propagates radially toward the stator mid blade height and then affects the upstream tandem rotor row. This study explains how a VRT blade mechanism with 8° of positive rotation of the aft rotor controls and suppresses the instability well before the stall zone affects the whole stage.

The inlet attack angle of rotor blade reasonably can be adjusted with the change of the stagger angle of inlet guide vane (IGV); so the efficiency of each condition will be affected. For the purpose to improve the efficiency, the DSP (Digital Signal Processor) controller is designed to adjust the stagger angle of IGV automatically in order to optimize the efficiency at any operating condition. The A/D signal collection includes inlet static pressure, outlet static pressure, outlet total pressure, rotor speed and torque signal, the efficiency can be calculated in the DSP, and the angle signal for the stepping motor which control the IGV will be sent out from the D/A. Experimental investigations are performed in a three-stage, low-speed axial compressor with variable inlet guide vanes. It is demonstrated that the DSP designed can well adjust the stagger angle of IGV online, the efficiency under different conditions can be optimized. This establishment of DSP online adjustment scheme may provide a practical solution for improving performance of multi-stage axial flow compressor when its operating condition is varied.

Chapter 7 - Turbomachinery

The aim of the paper is to explore the influence of the reversed blade-angle slot casing treatment (RBSCT) and its axial deflected angle (ADA) on the compressor performance and stability, and to reveal the mechanism that the change in ADA of the RBSCT influences the effect to broaden the compressor stable working range. The NASA Rotor 35 is used as the object of the investigation, and four RBSCTs with ADA of −15°, −30°, −45° and −60° are designed and investigated by unsteady numerical simulation. The results show that as the absolute value of the axial deflected angle increases, the capacity to improve the compressor stability of the RBSCT increases and then decreases. The unsteadiness of the injection and suction flows formed by the reversed blade angle slot plays an important role in the removal of the low-velocity zone. When ADA is −30°, the unsteadiness amplitude of the injection and suction flows is significantly higher than those of the other three. Consequently, the RBSCT with −30° ADA obtains the maximum stall margin improvement of 17.41% and the maximum design point efficiency improvement of 1.06% among the four RBSCTs.

Over the years it has been speculated that the performance of multi-stage axial flow compressors is enhanced by the passage of a wake through a blade row prior to being mixed-out by viscous diffusion. The link between wake mixing and performance depends on the ability to recover the total pressure deficit of a wake by a reversible flow process. This paper shows that such a process exists, it is unsteady, and is associated with the kinematics of the wake vorticity field. The analysis shows that the benefits of wake total pressure recovery can be estimated from linear theory and quantified in terms of a volume integral involving the deterministic stress and the mean strain rate. In the limit of large reduced frequency the recovery process is shown to be a direct function of blade circulation. Results are presented which show that the recovery process can reduce the wake mixing loss by as much as seventy percent. Under certain circumstances this can lead to nearly a point improvement in stage efficiency, a nontrivial amount.

Design factors which influence rotating stall recoverability of a high-speed multistage compressor are not yet fully understood. A high-speed, 10-stage compressor component was tested while operating in-stall to investigate parameters that affect the overall recoverability of a multistage compressor. The compressor instrumentation and data acquisition procedures were designed to obtain detailed performance data from the compressor while it entered into a rotating stall condition and while it operated in rotating stall. The compressor was tested at different in-stall operating conditions by varying compressor shaft speed, discharge throttle, and variable geometry settings to determine the effect of each variable on rotating-stall performance and recoverability. Test results suggest that the stall cell may not extend the full length of the compressor but instead can be confined to a portion (here the rear stages) of the compressor. When stages are stacked together, as is the case in a multistage compressor, the stalled performance and subsequent recoverability are greatly affected by how well the stages of the compressor are matched. The results also suggest that high-speed flows in the tenth stage may extend in-stall operation causing low recoverability of the overall test compressor at higher shaft speeds.

A bend skewed casing treatment was designed, to study the influence of one of its geometrical parameter porosity on the stable performance of single stage transonic axial flow compressor. The compressor was designed for the stage total-to-total pressure ratio of 1.35, corrected mass flow rate of 22 kg/s at corrected design speed of 12930 RPM. Bend skewed casing treatment has an axial inlet segment till 50% of the total length and rear segment that is skewed by 45° in the direction of the rotor tip section stagger. Both the sections were oriented at a skew angle of 45° to the radial plane such that the flow exiting the slot is in counter-clockwise direction to that of the rotor direction. The casing treatment slot width was equal to the maximum thickness of the rotor blades. Three casing treatment configurations were identified for the current experimental investigation. All the treatment geometries considered for the experimental research have lower porosities than reported in the open literatures. The effect of the porosity parameter on the performance of transonic compressor stage was evaluated at two axial coverages of 20% and 40% relative to the rotor tip axial chord. Performance maps were obtained for the solid casing and casing treatment with three different porosities. Comparative studies were carried out and experimental results showed a maximum of 65% improvement in the stable operating range of the compressor for one of the treatment configurations. It was also observed that the stable operating range of the compressor increases with an increase in the casing treatment porosity. All the casing treatment configurations showed that the compressor stall occurs at lower mass flows as compared to the solid casing. Compressor stage peak efficiency shows significant degradations with increase in the porosity as compared to solid casing. Detailed blade element performances were also obtained using calibrated multi-hole aerodynamic probe. Comparative variations of flow parameters like absolute flow angle, Mach number were studied at full flow and near stall conditions for the solid casing and casing treatment configurations. Hot wire measurements show very high fluctuation in the inlet axial velocity in the presence of solid casing as compared to casing treatments. Experimental investigation revealed that the porosity of the casing treatments has strong influence on the transonic compressor stage performance.

The internal losses in the tip clearance region strongly influence the compressor performance and its operational range. Previous research proved that passive wall treatments with circumferential grooves in axial compressors effectively increase the compressor stall margin. The vortex generated inside the circumferential grooves creates a resistance to the flow that leaks into the tip clearance region of the compressor. However, most works found in the literature on circumferential grooves in axial compressors deal only with high-performance single-stage axial compressors. Therefore, there is a need to investigate and analyze the behavior of circumferential grooves in a multi-stage environment. In the present work, a passive wall treatment with circumferential grooves was implemented in a multi-stage axial compressor. Different configurations of circumferential grooves were created at the casing of the first and second rotor rows used in a four-stage axial flow compressor. Numerical simulations were performed to evaluate the influence of the circumferential grooves on the performance of a multi-stage axial compressor. The results obtained after the simulations for the different circumferential groove configurations were compared with the results obtained for the compressor without casing treatment (smooth wall) for different rotational speeds. Furthermore, the complete compressor map characteristics were simulated for the different casing treatment configurations, and the results were compared with the compressor characteristics of the smooth wall case. The passive wall treatment with circumferential grooves produced changes in the multi-stage axial compressor flow field, especially in the tip clearance region, improving the compressor stability mainly for part load speeds.

Numerical simulations were carried out with the purpose of investigating the effect of applying circumferential grooves at axial compressor casing passive wall treatment to enhance the stall margin and change the tip leakage flow. The tip leakage flow is pointed out as one of the main contributors to stall inception in axial compressors. Hence, it is of major importance to treat appropriately the flow in this region. Circumferential grooves have shown a good performance in enhancing the stall margin in previous researches by changing the flow path in the tip clearance region. In this work, a passive wall treatment with four circumferential grooves was applied in the transonic axial compressor NASA Rotor 37. Its effect on the axial compressor performance and the flow in the tip clearance region was analyzed and set against the results attained for the smooth wall case. A 2.63% increase in the operational range of the axial compressor running at 100%N, was achieved, when compared with the original smooth wall casing configuration. The grooves installed at compressor casing, causes an increase in the flow entropy generation due to the high viscous effects in this gap region, between the rotor tip surface and casing with grooves. These viscous effects cause a drop in the turbomachine efficiency. For the grooves configurations used in this work, an efficiency drop of 0.7% was observed, compared with the original smooth wall. All the simulations were performed based on 3D turbulent flow calculations using Reynolds Averaged Navier-Stokes equations, and the flow eddy viscosity was determined using the two-equation SST turbulence model. The details of the grooves geometrical dimensions and its implementation are described in the paper.

Tip clearance has great influence on the performance of multistage axial compressors including efficiency, pressure rise, mass flow, as well as matching. This paper reports a study into the influence of tip clearance on the performance and matching of a 5-stage axial compressor by a numerical method. Different tip clearances from 0% to 5.0% span which represents the typical range of tip clearance in modern multistage axial compressors were simulated and analyzed. The results show that as tip clearance increases from 0% to 5.0% span, the choked mass flow decreases by about 21.8%, the peak pressure ratio decreases by about 43.1% and the peak efficiency decreases about 14.3 percents. As tip clearance increases, the efficiency of the whole compressor decreases in a parabolic manner not linearly as previous suggested, which is partially attributable to the cantilevered stators considered in this paper and primarily due to the mismatching of different stages. It is of great importance to control the tip clearance. When tip clearance increases, the front stage tends to work near surge condition and the rear stage tends to work near choke condition, which leads to lower efficiency than in the middle stages. A weight was defined to evaluate each stage’s contribution to the whole compressor’s efficiency deficit caused by the increase of tip clearance. Front and rear stages contribute more to the efficiency deficit than the middle stages, which indicates that more attention should be paid on front and rear stages to improve the performance of multistage axial compressors. In order to evaluate the matching of multistage axial compressors with a quantified method, a new parameter named “Peak Efficiency Deviation (PED)” was defined based on the difference between each stage’s operating efficiency and its peak efficiency. The mass flow of multistage axial compressors should be well considered to make the PED parameter to be close to zero as possible. In the most commonly used range of tip clearance from 0.5% to 3.0% span, the PED varies little within 0.4 percent, which is only about 8.4% of the peak efficiency deficit at 1.5% span tip clearance. So, the PED could be small within a wide range of tip clearances if the matching of the compressor is perfect at design tip clearance.

Computational analyses are carried out to predict the effect of the hub leakage flow from the inner banded stator cavities on the overall performance of an axial high pressure compressor. The results of a full fidelity simulation model, which includes cavities, sealing and main flow, is compared with models without the hub leakage flow. The CFD prediction confirms a significant effect from the hub leakage on the performance of the multistage axial compressor. A simpler and faster CFD modeling technique is explored and discussed for the modeling of the effect of the hub cavity leakage. In this approach, costly details of the stator cavities are ignored. Transfer functions or correlations are derived from the 1.5 stage (IGV-Rotor1-Stator1) of a multistage compressor having inner banded stator cavity. These correlations are used as boundary conditions to the primary flow path of the compressor while carrying out the simplified CFD models for the complete 4 stage compressor model. A MINITAB study is conducted to understand the most influencing parameters and their interactions in deriving the correlations. Analyses are then carried out using the simplified model with the hub leakage boundary conditions derived from these correlations. The CFD results of the full fidelity simulation and the simplified models are compared in this paper.

Variability in stator vanes is a widely used technique to improve the stability and efficiency of axial flow compressor in gas turbine engines. Most of the modern aircraft jet engines use variable stator vanes in both low pressure and high pressure compressors primarily for off-design performance. This study discusses in detail about the effect of stator variability in a three stage low pressure axial compressor at design and off-design conditions. Computational flow analysis were carried out for the three stage low pressure compressor with variability in inlet guide vane and first stage stator blade. Detailed investigation on flow physics was carried out in rotor blade passages with stator variability. At off-design speeds, the reduction in flow velocity is lower than the reduction in blade tip speed. This leads to mismatch in flow angles and inlet blade angles causing high incidence and large flow separation in blade passage. This results in poor aerodynamic stability of the axial compressor at off-design speeds. In this study, aerodynamic performance of compressor is evaluated from 70% to 100% design speeds with different stagger angle setting of inlet guide vane at each speed. Further, to improve 2nd stage rotor performance, variability was introduced in 1st stage stator blade and performance was evaluated. Compressor test results are compared with CFD data for design and off-design speeds.

A five-stage axial-flow compressor with all rotors operating with transonic relative inlet Mach numbers was designed as a research vehicle at the Lewis Research Center in 1952. The compressor was designed and tested as a component of a turbojet engine. This paper summarizes the research work done on this compressor including the aerodynamic design and detailed performance characteristics.

Sweep is often applied to transonic rotors to reduce the shock loss and to stator vanes to reduce the end-wall separation loss. In this paper, a new type of sweep is created for cantilevered stators in a 5-stage axial-flow compressor through a combined use of sensitivity analysis and an integrated optimization design method. The shape of sweep is focused on the stator hub region according to sensitivity analysis, and the new swept stator design is beneficial for reducing the stator loss at hub end and improving the compressor aerodynamic performance. Firstly, the optimization design of a single first stator row is conducted in the multistage environment, and the aerodynamic benefits gain is mainly due to two aspects. On the one hand, the effect of sweep near the hub reduces the stator inlet Mach number and further reduces the size of supersonic regions and the peak Mach number on the suction surface, thus decreasing the aerodynamic losses. On the other hand, the sweep also reduces the loading in the front part of the blade near the hub and weakens the leakage flow, thus reducing the flow blockage and decreasing the secondary losses in the hub region. After the optimization, the stator loss is reduced by 7.4%, leading to an increment of 0.21% of adiabatic efficiency of the multistage compressor. Secondly, the same optimization strategy is then applied to the remaining stator rows to maximize the aerodynamic benefits gain. The optimized compressor achieves an approximately 1% improvement of adiabatic efficiency over than the baseline one and also performs better over the whole operating range. Simultaneously, the operating range of the multistage compressor is widened. The results indicate that cantilevered stators with the type of sweep in this work have a significant potential to improve the performance of multistage axial-flow compressor.