Junqiang Zhu

Effects of low Reynolds number on flow stability of a transonic compressor

As the aircraft cruising at high altitude over 20,000 m with subsonic speed, the Reynolds number in terms of the compressor blade becomes very low and the compressor performance decreases dramatically due to separation of boundary layer and secondary-flow. The main objective in this paper is to understand the physical mechanism by which Reynolds number affects the compressor stable range. In this paper, a series of steady and unsteady numerical simulations were carried out for a transonic compressor rotor under several conditions, which corresponded to the operations at sea level, and at high altitude. Detailed analyses of the flow visualization have exposed the different flow topologies of the complicated secondary flow. It was found that the transonic axial-flow compressor rotor used in current investigation was prone to tip stall behavior, and the complex flow mechanisms which occur near the blade tip are found to be the key factors for the limited flow stability both under high Reynolds number and low Reynolds number. Under high Reynolds number, the tip flow field was dominated by the low momentum zones generated by the interaction between tip leakage flow and incoming flow, and with the decrease of Reynolds number, the tip leakage flow was weakened and the low-momentum fluid due to the surface boundary layer separation and radial transport of low-energy fluid was dominant in the tip flow.

Mechanism of the Interaction Between Casing Treatment and Tip Leakage Flow in a Subsonic Axial Compressor

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.Copyright

Parametric Studies of Pipe Diffuser on Performance of a Highly Loaded Centrifugal Compressor

Pipe diffusers with several different geometries were designed for a highly loaded centrifugal compressor originally using a wedge diffuser. Parametric studies on the effect of pipe diffuser performance of a highly loaded centrifugal compressor by varying pipe diffuser inlet-to-impeller exit radius ratio, throat length, divergence angle, and throat area on centrifugal compressor performance were performed using a state-of-the-art multi-block flow solver. An optimum design of pipe diffuser was obtained from the parametric study, and the numerical results indicate that this pipe diffuser has remarkable advantageous effects on the compressor performance. Furthermore, a detailed comparison of flow visualization between the pipe diffuser and the wedge diffuser was conducted to identify the physical mechanism that account for the beneficial effects of the pipe diffuser on the performance and stability of the compressor. It was found that the performance enhancement afforded by the pipe diffuser is a result of the unique diffuse inlet flow pattern. Alleviating flow distortion in the diffuser inlet and reducing the possibility of a flow separation in discrete passages are the physical mechanisms responsible for improving the highly loaded centrifugal compressor performance.

Effect of Fuel Staged Proportion on NOX Emission Performance of Centrally Staged Combustor

The low NOX emission technology has become an important feature of advanced aviation engine. A wide range of applications attempt to take advantage of the fact that staged combustion operating under lean-premixed-prevaporized (LPP) conditions can significantly decrease pollution emissions and improve combustion efficiency. In this paper a scheme with fuel centrally staged and multi-point injection is proposed. The mixing of fuel and air is improved, and the flame temperature is typically low in combustion zone, minimizing the formation of nitrogen oxides (NOX), especially thermal NOX. In terms of the field distribution of equivalence ratio and temperature obtained from Computational Fluid Dynamics (CFD), a chemical reactor network (CRN), including several different ideal reactor, namely perfectly stirred reactor (PSR) and plug flow reactor (PFR), is constructed to simulate the combustion process. The influences of the pilot equivalence ratio and percentage of pilot/main fuel on NOX and carbon monoxide (CO) emissions were studied by Chemical CRN model. Then the NOX emission in the staged combustor was researched experimentally. The effects of the amount of pilot fuel and primary fuel on pollution emissions were obtained by using gas analyzer. Finally, the effects of pilot fuel proportion on NOX emission were discussed in detail by comparing predicts of CRN and experimental results.Copyright

Experimental and Numerical Investigation of a Subsonic Compressor With Bend Skewed Slot Casing Treatment

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.© 2006 ASME

Investigation for the Effects of Circumferential Grooves on the Unsteadiness of Tip Clearance Flow to Enhance Compressor Flow Instability

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.Copyright

Study of Flow Control Using a Slotted Blade for a Compressor Airfoil at Low Reynolds Numbers

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 1.2×105 to 2.1×105 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.Copyright

Behavior of Tip Leakage Flow in an Axial Flow Compressor Rotor

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.Copyright

Effects of periodic wakes on boundary layer development on an ultra-high-lift low pressure turbine airfoil

The laminar-turbulent transition process in the boundary layer is of significant practical interest because the behavior of this boundary layer largely determines the overall efficiency of a low pressure turbine. This article presents complementary experimental and computational studies of the boundary layer development on an ultra-high-lift low pressure turbine airfoil under periodically unsteady incoming flow conditions. Particular emphasis is placed on the influence of the periodic wake on the laminar-turbulent transition process on the blade suction surface. The measurements were distinctive in that a closely spaced array of hot-film sensors allowed a very detailed examination of the suction surface boundary layer behavior. Measurements were made in a low-speed linear cascade facility at a freestream turbulence intensity level of 1.5%, a reduced frequency of 1.28, a flow coefficient of 0.70, and Reynolds numbers of 50,000 and 100,000, based on the cascade inlet velocity and the airfoil axial chord length. Experimental data were supplemented with numerical predictions from a commercially available Computational Fluid Dynamics code. The wake had a significant influence on the boundary layer of the ultra-high-lift low pressure turbine blade. Both the wake’s high turbulence and the negative jet behavior of the wake dominated the interaction between the unsteady wake and the separated boundary layer on the suction surface of the ultra-high-lift low pressure turbine airfoil. The upstream unsteady wake segments convecting through the blade passage behaved as a negative jet, with the highest turbulence occurring above the suction surface around the wake center. Transition of the unsteady boundary layer on the blade suction surface was initiated by the wake turbulence. The incoming wakes promoted transition onset upstream, which led to a periodic suppression of the separation bubble. The loss reduction was a compromise between the positive effect of the separation reduction and the negative effect of the larger turbulent-wetted area after reattachment due to the earlier boundary layer transition caused by the unsteady wakes. It appeared that the successful application of ultra-high-lift low pressure turbine blades required additional loss reduction mechanisms other than “simple” wake-blade interaction.

Numerical research on the mixing mechanism of lobed mixer with inlet swirl in linear radial distribution

A detailed numerical simulation is presented to investigate the effects of inlet swirl and its radial distribution on the mixing mechanisms of a turbofan mixer with 12 lobes, by using the commercial ANSYS CFX solver and k–ω SST model. The core-to-bypass temperature ratio and pressure ratio were set to 2.59 and 0.97, respectively, giving the Mach number of 0.66 and bypass ratio of 2.65 at mixing nozzle outlet. In the core inlet, the swirl angle was raised from 0° to 30° in a uniform or linear radial distribution manner. The inlet swirl and its radial gradient did enhance the development, interaction, and dissipation of the vortices downstream of lobed mixer, resulting in accelerating the lobed jet mixing. When the inlet swirl was less than 20°, the total pressure and thrust loss increments of lobed jet were acceptable and no more than 0.26% and 1.57%, respectively, compared with the baseline case. The results also showed that the three-dimensional separation bubble on center-body and the backflows along jet axis at the rail of center-body, resulting from the swirling flow between lobes’ trough and center-body, were the dominant sources of total pressure and thrust losses for all cases with inlet swirl. And, reasonable radial distribution of inlet swirl could inhibit the aforementioned 3D separation and backflow, and thus limited the increment of jet mixing loss favorably.