As a degree of freedom in the three-dimensional blade design of axial compressors, the sweep technique significantly affects the aerodynamic performance of axial compressors. In this paper, the effects of backward sweep rotor configurations on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at different rotational design speeds are studied by numerical simulation. The aim of this work is to improve understanding of the flow mechanism of backward sweep on the aerodynamic performance of a highly loaded axial compressor. A commercial CFD package is employed for flow simulations and analysis. The study found that at the design rotational speed, compared with baseline, backward sweep rotor configurations reduce the blade loading near the leading edge but slightly increases the blade loading near the trailing edge in the hub region. As the degree of backward sweep increases, the stall margin of the 1.5-stage axial compressor increase first and then decrease. Among different backward sweep rotor configurations, the 10% backward sweep rotor configuration has the highest stall margin, which is about 2.5% higher than that of baseline. This is due to the change of downstream stator incidence, which improves flow capacity near the hub region. At 80% rotational design speed, backward sweep rotor configurations improve stall margin and total pressure ratio of the compressor. It’s mainly due to the decreases of the rotor incidence near the middle span, which results in the decreases of separation on the suction surface. At 60% rotational design speed, detached shock disappears. Backward sweep rotor configurations deteriorate stall margin of the compressor, but increase total pressure ratio and adiabatic efficiency when the flow rate is lower than that at peak efficiency condition. Therefore, it’s necessary to consider the flow field structure of axial compressors at whole operating conditions in the design process and use the design freedom of sweep to improve the aerodynamic performance.

The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Centrifugal compressor stages with pipe diffusers are characterized by their high efficiency, especially under high pressure ratio conditions. Although it is believed that pipe diffuser scallop leading edge formed by the intersection of two pipes is a critical point in pipe diffuser design, there is another crucial and influential point, which is how to guide and decelerate the flow from pipe diffuser throat to the inlet of combustor chamber, with minimum loss and maximum outflow uniformity. Fishtail pipe diffuser passage is employed by Pratt&Whitney to connect impeller exit and combustor chamber inlet due to its improved performance characteristics. However, only a few comprehensive results have been published describing the complex flow patterns in the fishtail diffuser. Therefore, in the present work fishtail pipe diffusers with several different geometries were designed for a pressure ratio 8.3 centrifugal compressor stage used on a small turbo engine, aiming at providing detailed understanding of geometric parameter influence on fishtail pipe diffuser performance and flow mechanisms in complex fishtail passages. Cone length, streamwise area distribution and centerline shape are critical parameters of a fishtail pipe diffuser. Hence, parametric studies on fishtail pipe diffuser of this high pressure ratio centrifugal compressor by varying cone length, area distribution and centerline shape of the diffuser passage were performed using a state-of-the-art multi-block flow solver. These three parameters were changed respectively, while keeping other parameters unchanged. Detailed analysis was done to identify the influence on flow field in fishtail diffuser passage when these parameters were changed. It was found that increase of fishtail diffuser cone length could alleviate separation in diffuser passage, thus compressor performance is improved. And linear area distribution along passage centerline could build a more efficient fishtail pipe diffuser. A trumpet-shaped or bell-shaped passage is more likely to make flow separate. The centerline is of vital importance for a fishtail passage and it was built by two lines tangent to an ellipse in this work. It was modified by changing major and minor axes of the ellipse. Stage total pressure ratio and adiabatic efficiency maps for varying fishtail passage centerlines were obtained by numerical method, which indicate that there is an optimum range for both axes to make the fishtail pipe diffuser have a better performance. Through these works, the geometric parameter influence on fishtail diffuser performance was uncovered and physical insight into complex flows in fishtail pipe diffuser passage was obtained to give some guidelines on diffusing system design with fishtail pipe diffuser.

This present work is aimed at providing detailed understanding of the flow mechanisms in a highly loaded centrifugal compressor with different diffusers. Performance comparison between compressor stages with pipe diffuser and its original wedge diffuser was conducted by a validated state-of-the-art multi-block flow solver at different rotating speeds. Stage with pipe diffuser achieved a better performance above 80% rotating speed but a worse performance at lower rotating speeds near surge, than that of stage with wedge diffuser. Four operating points including the design point were analyzed in detail. The inherent diffuser leading edge of pipe diffuser could alleviate the flow distortion upstream diffuser throat and created a better operating condition for the downstream diffusion, which reduced the possibility of flow separation in discrete passages at design rotating speed. At 60% rotating speed operating point, there was a misalignment between the leading edge absolute flow angle and the metal angle of diffuser, resulted in an acceleration near diffuser leading edge due to the large negative incidence angle. The sharp leading edge of pipe diffuser could largely accommodate this negative incidence as comparison of the round leading edge of wedge diffuser. As a result, the flow separation was depressed and a better performance was achieved in the pipe diffuser. At 60% rotating speed near surge, performance of the pipe diffuser dropped below wedge diffuser. Total pressure loss of pipe diffuser exceeded that of the wedge diffuser due to the larger friction loss near wall at throat and cone, meanwhile ineffective static pressure recovery for pipe diffuser was triggered by the strong boundary layer blockage in the front of pipe diffuser cone.

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.

Flow control using a blade which is slotted from pressure surface to suction surface was studied. The effects of a slotted blade for a compressor airfoil at low Reynolds numbers were studied with the help of a linear cascade wind tunnel and a state-of-the-art multi-block flow solver. Measurements were taken in a low-speed linear cascade. The compressor cascade with a steady inflow was examined based on blade chord length at low Reynolds numbers. The experimental results showed that the slotted blade could effectively reduce the loss wake width and depth. The maximum relative reduction of wake width and depth were 48% and 45%, respectively. The numerical results at Reynolds numbers ranging from 1.2×10 5 to 2.1×10 5 showed that the airfoil aerodynamics of cascade are substantially improved by a slotted blade within this Reynolds number regime. This yields less total pressure loss and a greater flow turning angle and static pressure rise. The extensive numerical results showed that the slot inlet could capture the high pressure flow from the pressure surface, bring it into the slot, and discharge it onto the suction surface of the blade. This then blows off the boundary layer separation flow and accelerates the low energy flow, reducing the size of the separation region, or even eliminating the separation entirely on the trailing edge of blade.

The use of slots and grooves in the shroud over the tips of compressor blades, known as casing treatment, is a powerful method to control tip leakage flow through the clearance gap and enhance the flow stability in compressors. This paper presents a contribution to the understanding of the physical mechanism by which circumferential groove casing treatment manipulates the tip clearance flow’s unsteadiness. A series of computational studies were carried out to understand the physical mechanism responsible for improvement in stall margin of a high subsonic axial-flow compressor rotor due to the circumferential groove casing treatment from an unsteady viewpoint. Detailed analyses of the flow visualization at the tip have exposed the different tip flow topologies between the cases with circumferential groove and with untreated smooth wall. It was found that the primary stall margin enhancement afforded by the circumferential groove casing treatment is a result of the unsteady tip clearance flow manipulation. Breaking balance of incoming/tip clearance flow axial momentum by inducing the radial movement and tangential movement and delay the occurrence of tip clearance’s unsteadiness are the physical mechanisms responsible for extending the compressor stall margin.

The phenomenon of flow instability in the compression system such as fan and compressor has been a long-standing “bottle-neck” problem for gas turbines/aircraft engines. With a vision of providing a state-of-the-art understanding of the flow field in axial-flow compressor in the perspective of enhancing their stability using passive means. Two topics are covered in this paper. The first topic is the stability-limiting flow mechanism close to stall, which is the basic knowledge needed to manipulate end-wall flow behavior for the stability improvement. The physical process occurring when approaching stall and the role of complex tip flow mechanism on flow instability in current high subsonic axial compressor rotor has been assessed using single blade passage computations. The second topic is flow instability manipulation with casing treatment. In order to advance the understanding of the fundamental mechanisms of casing treatment and determine the change in the flow field by which casing treatment improve compressor stability, systematic studies of the coupled flow through a subsonic compressor rotor and various end-wall treatments were carried out using a state-of-the-art multi-block flow solver. The numerically obtained flow fields were interrogated to identify complicated flow phenomenon around and within the end-wall treatments and describe the interaction between the rotor tip flow and end-wall treatments. Detailed analyses of the flow visualization at the rotor tip have exposed the different tip flow topologies between the cases with treatment casing and with untreated smooth wall. It was found that the primary stall margin enhancement afforded by end-wall treatments is a result of the tip flow manipulation. Compared to the smooth wall case, the treated casing significantly dampen or absorb the blockage near the upstream part of the blade passage caused by the upstream movement of tip clearance flow and weakens the roll-up of the core vortex. These mechanisms prevent an early spillage of low momentum fluid into the adjacent blade passage and delay the onset of flow instability.

This paper investigates the flow field at the tip region of compressor rotor. In particular, the effect of stepped tip gaps on the performance and flowfield of an axial-flow compressor rotor was reviewed using both experimental and computational methods. An axial compressor rotor with no inlet guide vanes was tested under subsonic condition. A parametric study of clearance levels and step profiles was performed using eight different casing geometries. This study was aimed at comparing compressor performance in specified configurations. The experimental results showed that the inclusion of stepped tip gaps with the small clearance level gave increased pressure ratio, efficiency, and stall margin throughout the mass flow range at both speeds. However, when using medium and large clearance level, the benefits of stepped tip gaps were not noticed for all rotor operating conditions if compared with the baseline case. Steady-state Navier-Stokes analyses were performed for cases involving small clearance level and stepped tip gap geometries. They highlighted the mechanisms associated with performance improvement. The numerical procedure correctly predicted the overall effects of stepped tip gaps. Detailed numerical simulation results showed that the interaction between the stepped groove flow and blade passage flow could entrain the blockage produced by upstream tip leakage flow into the tip gap of adjacent blades of the compressor rotor. It is through this process that stepped tip gaps can help dissipating blockage that was caused by upstream tip leakage flow. Thus the path and extent of the blockage in the tip region is altered to increase the passage throughflow area and so, the rotor performance can be improved.

Based on the test results of discrete axial and blade angle slot casing treatment, a new type of casing treatment was designed for a subsonic axial flow compressor rotor by optimising various geometry parameters. To obtain a wide operating range and to minimize penalties in terms of isentropic efficiency, seven compressor configurations incorporating casing treatments of 0%, 16.6%, 33.3%, 50%, 66.6%, 83.3% and 100% rotor exposure were experimentally investigated. The results showed that significant improvements in stall margin are possible in all exposures and insignificant isentropic efficiency sacrifices are recorded in some exposures. Nearly 21.43% stall margin improvement in terms of the corrected mass flow rate was achieved with 33.3% rotor blade tip axial chord exposure. The compressor build with 16.6% rotor exposure was the best configuration in terms of maximum isentropic efficiency gain. The second issue of the paper was to offer a contribution to the understanding of the physical mechanism by which bend skewed slot casing treatment improve stall margin under subsonic conditions. By applying a concept similar to “Domain Scaling” approach (as often used in multistage turbomachinery Flow-fields) to the interface between the rotor blade passage and end-wall treatments, a time-dependent 3-dimentional numerical simulation was performed for the subsonic axial-flow compressor rotor with bend skewed slot casing treatment. The numerical results agreed well with experimental results. Detailed analyses of the coupled flow through bend skewed slot casing treatment and rotor blade passage under subsonic conditions led to some preliminary conclusions as to the flow physics involved in the stall margin improvements afforded by the use of bend skewed slot casing treatment.

The current paper reports on investigations with an aim to advance the understanding of the flow field near the casing of a small-scale high-speed axial flow compressor rotor. Steady three dimensional viscous flow calculations are applied to obtain flow fields at various operating conditions. To demonstrate the validity of the computation, the numerical results are first compared with available measured data. Then, the numerically obtained flow fields are analyzed to identify the behavior of tip leakage flow, and the mechanism of blockage generation arising from flow interactions between the tip clearance flow, the blade/casing wall boundary layers, and non-uniform main flow. The current investigation indicates that the “breakdown” of the tip leakage vortex occurs inside the rotor passage at the near stall condition. The vortex “breakdown” results in the low-energy fluid accumulating on the casing wall spreads out remarkably, which causes a sudden growth of the casing wall boundary layer having a large blockage effect. A low-velocity region develops along the tip clearance vortex at the near stall condition due to the vortex “breakdown”. As the mass flow rate is further decreased, this area builds up rapidly and moves upstream. This area prevents incoming flow from passing through the pressure side of the passage and forces the tip leakage flow to spill into the adjacent blade passage from the pressure side at the leading edge. It is found that the tip leakage flow exerts little influence on the development of the blade suction surface boundary layer even at the near stall condition.

The use of slots and grooves in the shroud over the tips of compressor blades, known as casing treatment, is known as a powerful method to control tip leakage flow through the clearance gap and enhance the flow stability in compressors. This paper present a detailed steady and unsteady numerical studies of the coupled flow through rotor blade passages and two different types of casing treatment for a modern subsonic axial-flow compressor rotor. Particular attention was given to examining the interaction between the tip leakage flow and the casing treatment. In order to validate the multi block model applied in the rotor blade end-wall region, the computational results for the modern subsonic compressor rotor both with and without casing treatment were correlated with available experimental test data for estimation of the global performance. Detailed analyses of the flow visualization at the tip have exposed the different tip flow topologies between the cases with casing treatment and with untreated smooth wall. It was found that the primary stall margin enhancement afforded by the casing treatment is a result of the tip clearance flow manipulation. The repositioning of the tip clearance vortex further towards the trailing edge of the blade passage and delaying the movement of incoming/tip clearance flow interface to the leading edge plane are the physical mechanisms responsible for extending the compressor stall margin.