Reduction Of Aerodynamic Loss Research Articles

Research on modeling method of non-axisymmetric endwall based on body-fitted coordinates

Due to the large bending of turbine blades, there is strong three-dimensional flow in cascade channel, the three-dimensional flow at the endwall of blades generates extremely complex heat transfer characteristics. The design concept of high heat load turbine blades, particularly comprehensive optimization of flow and heat transfer at the endwall, has gained increasing attention. In this paper, we have developed a three-dimensional endwall parameterized modeling platform based on Body-Fitting Coordinates and introduced a novel approach for streamwise direction and normal direction concerning endwall profile construction. The endwall modeling was optimized using Genetic Algorithm to obtain a modified one with low heat transfer level on the premise that the aerodynamic loss is not increased. The analysis indicates that the reduction of aerodynamic loss with the modified endwall is relatively marginal, whereas the heat transfer intensity on the endwall surface experiences a significant decrease. Additionally, employing an endwall model can effectively mitigate the lateral pressure gradient in the cascade channel, which exerts a pronounced influence on suppressing secondary vortex development in the cascade.

Numerical Investigations of a Non-Uniform Stator Dihedral Design Strategy for a Boundary Layer Ingestion (BLI) Fan

A distributed propulsion system has the advantage of saving 5–15% fuel burn through ingesting the fuselage boundary layer of an aircraft by fan or compressor. However, due to boundary layer ingestion (BLI), the fan stage will continuously operate under serious inlet distortion. This will lead to a circumferentially non-uniform flow separation distribution on the stator blade suction surface along the annulus, which significantly decreases the fan’s adiabatic efficiency. To solve this problem, a non-uniform stator dihedral design strategy has been developed to explore its potential of improving BLI fan performance. First, the stator full-annulus blade passages were divided into blade dihedral design regions and baseline design regions on the basis of the additional aerodynamic loss distributions caused by BLI inlet distortion. Then, to find the appropriate dihedral design parameters, the full-annulus BLI fan was discretized into several portions according to the rotor blade number and the dihedral design parameter investigations for dihedral depth and dihedral angle were conducted at the portion with the largest inflow distortion through a single-blade-passage computational model. The optimal combinational dihedral design parameter (dihedral depth 0.3, dihedral angle 6 deg) was applied to the blade passages with notable flow loss which were mainly located in the annulus positions from −120 to 60 degrees suffering from inlet distortion, while the blades in the low-loss annulus locations were unchanged. In this way, a non-uniform stator dihedral design scheme was achieved. In the end, the effectiveness of the non-uniform stator dihedral design was validated by analyzing the internal flow fields of the BLI fan. The results show that the stator dihedral design in distorted regions can increase the inlet axial velocity and reduce the aerodynamic load near the blade trailing edge, which are beneficial for suppressing the flow separations and reducing aerodynamic loss. Specifically, compared with the baseline design, the non-uniform stator dihedral design has achieved a reduction of aerodynamic loss of about 7.7%. The fan stage has presented an improvement of adiabatic efficiency of about 0.48% at the redesigned point without sacrificing the total pressure ratio. In the entire operating range, the redesigned fan has also shown a higher adiabatic efficiency than the baseline design with no reduction of the total pressure ratio, which provides a probable guideline for future BLI distortion-tolerant fan design.

Aerodynamic optimisation of a winglet-cavity tip in a high-pressure axial turbine cascade

Tip clearance flow between rotating blades and the stationary casing in high-pressure turbines is very complex and is one of the most important factors influencing turbine performance. The rotor with a winglet-cavity tip is often used as an effective method to improve the loss resulting from the tip clearance flow. In this study, an aerodynamic geometric optimisation of a winglet-cavity tip was carried out in a linear unshrouded high-pressure axial turbine cascade. For the purpose of shaping the efficient winglet geometry of the rotor tip, a novel parameterisation method has been introduced in the optimisation procedure based on the computational fluid dynamics simulation and analysis. The reliability of a commercial computational fluid dynamics code with different turbulence models was first validated by contrasting with the experimental results, and the numerical total pressure loss and flow angle using the Baseline k-omega Model (BSL κ-ω model) shows a better agreement with the test data. Geometric parameterisation of blade tips along the pressure side and suction side was adopted to optimise the tip clearance flow, and an optimal winglet-cavity tip was proven to achieve lower tip leakage mass flow rate and total pressure loss than the flat tip and cavity tip. Compared to the numerical results of flat tip and cavity tip, the optimised winglet-cavity design, with the winglet along the pressure side and suction side, had lower tip leakage mass flow rate and total pressure loss. It offered a 35.7% reduction in the change ratio [Formula: see text]. In addition, the optimised winglet along pressure side and suction side, respectively, by using the parameterisation method was studied for investigating the individual effect of the pressure-side winglet and suction-side winglet on the tip clearance flow. It was found that the suction-side extension of the optimal winglet resulted in a greater reduction of aerodynamic loss and leakage mass flow than the pressure-side extension of the optimal winglet. Moreover, with the analysis based on the tip flow pattern, the numerical results show that the pressure-side winglet reduced the contraction coefficient, and the suction-side winglet reduced the aerodynamic loss effectively by decreasing the driving pressure difference near the blade tips, the leakage flow velocity, and the interaction between the leakage flow and the main flow. Overall, a better aerodynamic performance can be obtained by adopting the pressure-side and suction-side winglet-cavity simultaneously.

Multi-Objective Optimization of a Row of Film Cooling Holes Using an Evolutionary Algorithm and Surrogate Modeling

Multi-objective shape optimization of a row of laidback fan-shaped film cooling holes has been performed using a hybrid multi-objective evolutionary approach in order to achieve an acceptable compromise between two competing objectives: the enhancement of film cooling effectiveness and the reduction of aerodynamic loss. In order to perform comprehensive optimization of a film cooling hole shape, the injection angle of the hole, lateral expansion angle of the diffuser, forward expansion angle of the hole, and pitch-to-hole diameter ratio are chosen as design variables. Forty experimental designs within the design spaces are selected using the Latin hypercube sampling method. The response surface approximation method is used to construct the surrogate using objective function values calculated at the experimental points using Reynolds-averaged Navier-Stokes analysis. The shear stress transport turbulence model is used as a turbulence closure. The optimization results are processed using the Pareto-optimal method. The Pareto-optimal solutions are obtained using a combination of a evolutionary algorithm and a local search method. The optimum designs are grouped using the k-means clustering technique, and the three optimal points selected in the Pareto-optimal solutions are evaluated by numerical analysis. The optimum designs give enhanced objective function values compared to the experimental designs.

Tip gap flow and aerodynamic loss generation in a turbine cascade equipped with suction-side winglets

The tip gap flow and aerodynamic loss generation over a plane tip equipped with a “constant-width suction-side” (CWSS) winglet and a “varying-width suction-side” (VWSS) winglet have been investigated in a turbine cascade. For a fixed tip gap of h/c = 2.0%, three different winglet widths of w/p = 5.28, 10.55, and 15.83% are tested for the CWSS winglet. The VWSS winglet is designed based on flow visualization and has almost the same winglet area as the CWSS winglet of w/p = 15.83%. In general, the suction-side winglets have a role to increase aerodynamic loss in the tip leakage vortex region but reduce aerodynamic loss in the passage vortex region. For the CWSS winglet, the total-pressure loss coefficient mass-averaged all over the measurement plane has no appreciable changes with increasing w/p from 0.0 to 10.55%, but tends to decrease with further increment of w/p. The VWSS winglet performs better in reducing aerodynamic loss in the passage vortex region than the CWSS winglet of w/p = 15.83% but leads to a little bit higher aerodynamic loss in the tip leakage vortex region. The aerodynamic loss reduction by the VWSS winglet is 7.4% in comparison with the plane tip without winglet, and is about 60% lower than that by the widest CWSS winglet.

21004 超高負荷タービン円環翼列の空力性能評価(GS-20【流体機械】)

In this study, the aerodynamic performances of the three types of ultra-highly loaded turbine cascades, which are different in length along the suction surface (SS), were assessed by using the small annular turbine cascade test rig. The measurements for the torque and the total pressures at the inlet and the outlet of turbine stage were performed at same corrected mass flow rate in order to evaluate the turbine efficiency and the aerodynamic loss. The experimental results showed that the increase of the length of SS from the standard blade profile reduced the torque, but increased the turbine efficiency by the reduction of aerodynamic loss. On the other hand, the decrease of the length of SS more increased the reduction of torque than that in the increase of the length of SS, and consequently reduced the turbine efficiency.

터빈 동익 컷백스퀼러팁 하류에서의 3차원 유동 및 압력손실

The effect of channel cutback on three-dimensional flow fields and aerodynamic losses downstream of a cavity squealer tip has been investigated in a turbine rotor cascade for the squealer rim height-to-chord ratio and tip gap height-to-chord ratio of h<SUB>st</SUB>/c = 5.51% and h/c = 2.0% respectively. The cutback length-to-camber ratio is changed to be CB/c<SUB>c</SUB> = 0.0, 0.1, 0.2 and 0.3. The results show that longer cutback delivers not only stronger secondary flow but also higher aerodynamic loss in the tip leakage vortex region, meanwhile it leads to lower aerodynamic loss in the passage vortex region. The discharge of cavity fluid through the cutback opening provides a beneficial effect in the reduction of aerodynamic loss, whereas there also exists a side effect of aerodynamic loss increase due to local wider tip gap near the trailing edge. With increasing CB/c<SUB>c</SUB> from 0.0 to 0.3, the aerodynamic loss coefficient mass-averaged all over the measurement plane tends to increase slightly.

Tip gap height effects on the aerodynamic performance of a cavity squealer tip in a turbine cascade in comparison with plane tip results: part 2—aerodynamic losses

Tip gap height effects on aerodynamic losses downstream of a cavity squealer tip have been investigated in a linear turbine cascade for power generation, in comparison with plane tip results. Three-dimensional flow fields are measured with a five-hole probe for tip gap height-to-chord ratios of h/c = 0.5, 1.0, 1.5 and 2.0%. The cavity squealer tip has a full length squealer with its rim height-to-chord ratio of 5.51%. For a fixed value of h/c, the tip leakage vortex for the cavity squealer tip is always weaker than that for the plane tip, and the flow field in the passage vortex region for the cavity squealer tip is less influenced by the tip leakage flow than that for the plane tip. For the cavity squealer tip, there is no appreciable change in local aerodynamic loss with h/c in the passage vortex region, but local aerodynamic loss in the tip leakage vortex region increases with h/c. The roles of the cavity squealer tip in reducing aerodynamic loss in comparison with the plane tip case are twofold: (1) the cavity squealer tip decreases the leakage flow discharge in the region from the leading edge to the mid-chord, which leads to an aerodynamic loss reduction in the passage vortex region and (2) it also decreases the leakage flow discharge downstream of the mid-chord, which results in an aerodynamic loss reduction in the tip leakage vortex region.

Effects of squealer rim height on aerodynamic losses downstream of a high-turning turbine rotor blade

The effects of squealer rim height on three-dimensional flows and aerodynamic losses downstream of a high-turning turbine rotor blade have been investigated for a typical tip gap-to-chord ratio of h/ c = 2.0%. The squealer rim height-to-chord ratio is changed to be h st/ c = 0.00 (plane tip), 1.37, 2.75, 5.51, and 8.26%. Results show that as h st/ c increases, the tip leakage vortex tends to be weakened and the interaction between the tip leakage vortex and the passage vortex becomes less severe. The squealer rim height plays an important role in the reduction of aerodynamic loss when h st/ c ⩽ 2.75%. In the case of h st/ c ⩾ 5.51%, higher squealer rim cannot provide an effective reduction in aerodynamic loss. The aerodynamic loss reduction by increasing h st/ c is limited only to the near-tip region within a quarter of the span from the casing wall.