Qingjun Zhao

Multi-objective optimization of groove casing treatment in a transonic compressor

The paper presents a multi-objective optimization of circumferential casing grooves geometries for the NASA Rotor 37 transonic compressor. The depth normalized by the tip clearance and the width normalized by the tip chord are selected as the design variables. The stall margin and peak efficiency are used as the objective functions. The Latin Hypercube Sampling technique was used to select the sample points in the design space. Based on the numerical results of the sample points, the radial basis function network model of the artificial neural network was constructed. The NSGA-II multi-objective evolutionary algorithm is then employed to search for Pareto-optimal solutions. The leave-one-out cross validation method was also used to evaluate the precision of the radial basis function network model. The results of the optimization show the present method can be effectively used for the design of circumferential casing grooves to take account of the stall margin and efficiency. From the Pareto-optimal solutions, two groove configurations are selected and the internal flow fields are compared with the smooth casing. The effect mechanism of the circumferential casing grooves on the performance of the transonic compressor is discussed by the analysis of the flow in the blade tip region.

Shock loss model and blade profile optimization design of a supersonic cascade

With the improvement of single-stage compressor load, the inflow velocity and relative Mach number at blade tip of compressor are further increased. Using supersonic blade profile to design highly loaded compressor is an effective method to satisfy the requirements of the compressor aerodynamic design. In the internal flow of a blade row, detached shock is caused at the leading edge of supersonic cascade, and complex passage shock structure is produced by shock and boundary layer interaction at blade surface. Establishing shock loss calculation method for detached shock and passage shock and optimizing the blade profile according to the shock loss mechanisms are effective means to design high-efficiency compressors. The earlier studies of shock loss calculation model in supersonic compressor cascade have shown that shock loss of the detached shock at leading edge is related to leading edge radius and inlet Mach number. Therefore, several shock loss models for detached shock are established and proved to be effective. In previous studies on passage shock loss, the changes in geometry of passage shock, which is caused by shock and boundary layer interaction, are not fully considered. So, it is necessary to do further study on passage shock loss model. Based on the internal flow characteristics in supersonic cascade, passage shock loss model is proposed. Physics-based passage shock geometry, which is determined by static pressure rise, Mach number before wave fronts and blade profile curvature, is applied for this passage shock loss model. Near the blade surface, passage shock geometry is transformed into lambda-type shock that is composed of two branch oblique shocks, and oblique shock relationships are used to calculate their total pressure losses. Away from the blade surface, passage shock geometry is very close to the normal shock shape, so that normal shock relationships are used for loss calculation.

Tip-Clearance Effects on Hot-Streak Migration in Low-Pressure Stage of Vaneless Counter-Rotating Turbine

Three-dimensional multiblade row unsteady Navier-Stokes simulations have been performed to reveal the effects of rotor tip clearance on the inlet hot-streak migration characteristics in the low-pressure stage of a vaneless counter-rotating turbine. The numerical results indicate that most of the hotter fluid migrates toward the rotor pressure surface and that only a small amount of hotter fluid migrates to the rotor suction surface when it convects into the low-pressure turbine rotor. The hotter fluid that migrated to the tip region of the high-pressure turbine rotor impinges on the leading edge of the low-pressure turbine rotor after it goes through the high-pressure turbine rotor. The migration of the hotter fluid directly results in a very high heat load at the leading edge of the low-pressure turbine rotor. The leakage flow in the rotor tip clearance tends to increase the low-pressure turbine rotor outlet temperature at the tip region.

Tip Clearance Effects on Inlet Hot Streaks Migration Characteristics in High Pressure Stage of a Vaneless Counter-Rotating Turbine

In this paper, three-dimensional multiblade row unsteady Navier-Stokes simulations at a hot streak temperature ratio of 2.0 have been performed to reveal the effects of rotor tip clearance on the inlet hot streak migration characteristics in high pressure stage of a Vaneless Counter-Rotating Turbine. The hot streak is circular in shape with a diameter equal to 25% of the high pressure turbine stator span. The hot streak center is located at 50% of the span and the leading edge of the high pressure turbine stator. The tip clearance size studied in this paper is 2.0mm (2.594% high pressure turbine rotor height). The numerical results indicate that the hot streak mixes with the high pressure turbine stator wake and convects towards the high pressure turbine rotor blade surface. Most of hotter fluid migrates to the pressure surface of the high pressure turbine rotor. Only a few of hotter fluid rounds the leading edge of the high pressure turbine rotor and migrates to the suction surface. The migration characteristics of the hot streak in the high pressure turbine rotor are dominated by the combined effects of secondary flow, buoyancy and leakage flow in the rotor tip clearance. The leakage flow trends to drive the hotter fluid towards the blade tip on the pressure surface and to the hub on the suction surface. Under the effect of the leakage flow, even partial hotter fluid near the pressure surface is also driven to the rotor suction surface through the tip clearance. Compared with the case without rotor tip clearance, the heat load of the high pressure turbine rotor is intensified due to the effects of the leakage flow. And the results indicate that the leakage flow effects trend to increase the low pressure turbine rotor inlet temperature at the tip region. The air flow with higher temperature at the tip region of the low pressure turbine rotor inlet will affect the flow and heat transfer characteristics in the downstream low pressure turbine.Copyright

Aerodynamic performance analysis and optimization of a turbine duct with low degree of partial admission

A partial admission turbine duct with outlet-to-inlet area ratio greater than unity can increase the admission degree of the downstream turbine stage and, thus improve the performance of a multistage turbine with a low partial admission degree. However, the upstream flow structures of ducts, such as secondary flow, especially the circumferential nonuniformities originating from the effect of the partial admission, make the flow in ducts complex. The complexity of the flow has a negative impact on the performance of ducts. In the present investigation, numerical study of the flow behavior within ducts is done to evaluate the effect of the partial admission on the performance of the ducts. The study is carried out with regard to two cases, i.e. which are with the same duct geometry but are at different working conditions to highlight the impact of partial admission on the performance of ducts. Case 1 is used as baseline. It is designed based on circumferential mass-averaged flow conditions at ducts inlet. It causes the circumferential nonuniformities originating from the partial admission to have no impact on the performance of case 1. Case 2, which considers partial admission, is compared with case 1 to know the impact of the partial admission on the performance of ducts, and to give guidelines to design a duct for the partial admission turbines. Since the duct inlet conditions is a result of the interaction between partial admission turbine and duct, a straightforward way to consider the effect of the partial admission is to simulate the flows in ducts and upstream turbines contemporaneously. Comparative results indicate that the mixing of main flow in the admitted channel and the windage fluid from the unadmitted channel occurs at the duct inlet close to the duct circumferential wall. The adverse pressure gradient of case 2 in that region becomes larger than that of case 1. As a result, the flow separates at that region deteriorating the performance of ducts. Based on the simulation results of the previous cases, case 2’s circumferential wall surface, which is along the gas swirling direction is shrunk to accelerate the flow and, thereby, overcome the adverse pressure gradient imposed by the effect of the partial admission. The results show that the separation is restrained and the decrease in total pressure loss is 52.9%.

Effects of incidence angle on a low-pressure turbine blade boundary layer evolution through large eddy simulation:

The evolution mechanism of the boundary layer and coherent structures in a low-pressure turbine blade is discussed. Five different incidence angles over the T106A blade for a Mach number Ma = 0.404 and Reynolds number Re = 0.6 × 105 (based on the axial chord and outlet velocity) are performed using large eddy simulation method. The calculation results at +7.8 incidence angle are agreed well with the experimental and direct numerical simulation data. The influence of the incidence angle on the flow field is mainly shown at the front of the suction side and pressure side. As the incidence angle changes from positive to negative, the separation bubble near the leading edge disappears and the blade loading decreases gradually. When the incidence angle reduces to −5°, separation bubble appears near the leading edge of the pressure side. At the case of incidence angle equaling −10°, the length of time-averaged separation bubble on the pressure side grows to 39% axial chord and the evolution process of the coherent structures is extremely complex. The spanwise vortexes roll up near the leading edge and gradually evolve into streamwise vortexes. High-energy fluid in the main flow was driven to near-wall zone by the rotating effect of streamwise vortexes, which increases the fluid momentum inside the boundary layer. The streamwise vortexes are stretched by the strong acceleration of the flow until they transport to the trailing edge.

Investigation of groove casing treatment in a transonic compressor at different speeds with control volume method

This article uses a control volume method to analyze the circumferential groove casing treatment in a transonic compressor. To analyze the axial momentum transport through the tip gap, the control volume near the casing is divided into two parts: the control volumes inside and outside the tip gap. Besides, the association between the forces acting on the control volume and flow structures is studied by analyzing the distributions of axial momentum flux and axial shear stress. With this method, the flow mechanisms of stall margin improvement due to casing grooves in Rotor 35 are quantitatively analyzed. The analysis is conducted at 100% and 60% design speed with supersonic and subsonic tip speed, respectively. At design speed, the casing grooves decrease the axial shear force and the axial force due to the transport of axial momentum induced by the tip leakage flow. Meanwhile, the bleeding and injecting effect of grooves contribute much to the axial force due to the transport of axial momentum. Based on the axial distribution of the axial forces, the contribution of each groove to the stall margin improvement is assessed. And the grooves that play a major role in stall margin improvement are ascertained. At 60% design speed, because the blade loading is reduced, the axial momentum transport caused by the grooves cannot suppress the boundary layer separation effectively. Consequently, the stall margin of the compressor is not significantly improved by the casing grooves.

A new high-order unstructured numerical scheme for large eddy simulation

High dissipative second-order schemes are widely used in Reynolds Averaged NavierStokes (RANS) simulations owing to the robustness, while they are harmful to the numerical accuracy for large eddy simulation (LES). It is well known that high dissipative numerical schemes can be as influential as the sub-grid scale (SGS) models . In the present paper, a new high-order numerical scheme for unstructured grids is proposed by the authors. The scheme is based on a central-type reconstruction, which is non-dissipative. A circulative gradient correction is proposed to extend the stencils of reconstruction without a large CPU/memory requirement. To achieve high-order accuracy, the free parameter is determined by expending a Taylor expression. The dispersion property of the scheme is optimized by minimizing the numerical errors in large eddy simulations. Furthermore, a hybrid method combining the proposed scheme with a weighted essentially non-oscillatory (WENO) scheme is constructed to make it possible for the discontinuity-capturing. Several benchmark test problems have been solved to validate the properties of the present schemes. The results of a linear wave advection problem indicate that the numerical error of the present schemes is reduced by 1∼2 orders of magnitude compared with the traditional second-order schemes. The better dispersion properties are shown by solving the transportation of inviscid vortex, with no lagging or advancing phase error observed. The high resolution for turbulent flow is demonstrated by LES of turbulent channel flow. At last, the ability for shock-wave capturing of the hybrid WENO scheme is validated by a forward step problem, with a clear thrice reflecting structure present.

Unsteady Numerical Simulation of Shock Systems in Vaneless Counter-Rotating Turbine

A detailed unsteady numerical simulation has been carried out to investigate the shock systems in the high pressure (HP) turbine rotor and unsteady shock-wake interaction between coupled blade rows in a 1+1/2 counter-rotating turbine (VCRT). For the VCRT HP rotor, due to the convergent-divergent nozzle design, along almost all the span, fishtail shock systems appear after the trailing edge, where the pitch averaged relative Mach number is exceeding the value of 1.4 and up to 1.5 approximately (except the both endwalls). A group of pressure waves create from the suction surface after about 60% axial chord in the VCRT HP rotor, and those waves interact with the inner-extending shock (IES). IES first impinges on the next HP rotor suction surface and its echo wave is strong enough and cannot be neglected, then the echo wave interacts with the HP rotor wake. Strongly influenced by the HP rotor wake and LP rotor, the HP rotor outer-extending shock (OES) varies periodically when moving from one LP rotor leading edge to the next. In VCRT, the relative Mach numbers in front of IES and OES are not equal, and in front of IES, the maximum relative Mach number is more than 2.0, but in front of OES, the maximum relative Mach number is less than 1.9. Moreover, behind IES and OES, the flow is supersonic. Though the shocks are intensified in VCRT, the loss resulted in by the shocks is acceptable, and the HP rotor using convergent-divergent nozzle design can obtain major benefits.Copyright

The impact of circumferential casing grooves on rotating instability in a transonic axial compressor

The impact of circumferential casing grooves on rotating instability is first assessed for both design and part speed operations in a transonic axial compressor, with the purpose of developing the next generation casing treatments for vibration control. Multi-passage time-resolved computations are performed to capture the origination and propagation behavior of the instability for cases with and without casing grooves. Probed pressure signals in different passages show a nonsynchronous fluctuation of tip flow. It proves tip leakage vortex and its self-excited oscillation is responsible for this type of inconsistence, regardless of the compressor operation speed. Although flow separation on blade suction surface and the consequent shedding vortex contributes to another origin of instability, the resulted flow appears to be consistent. Casing grooves are able to enhance the synchronization by greatly suppressing both tip leakage vortex oscillations and the intermittently shedding separation vortex, especially in the front part of blade passage. Both types of instability are constrained in several separated axial scope by casing grooves, which essentially increase the damping of flow oscillations. Thus, further improvement of casing treatment design can be expected if the axial transport of the instability in the tip region is restrained more efficiently, for both extending stall margin and enhancing aerodynamic stability.