High-speed Compressor Cascade Research Articles

Experimental study on the effect of fluidic oscillator jet position on the performance of a high speed compressor cascade

In order to further improve the performance of the cascade, this paper focused on a high-load diffuser cascade and a pulsed-type fluidic oscillator for flow control was also applied. Experiments were conducted at design incidence of 0° and near stall incidence of +4°, investigating different jet positions of the fluidic oscillators. The results indicate that at Mach number of 0.67 and design incidence, the pulsed jet effect is optimal when the jet position of the fluidic oscillator is at 35% of chord, resulting in an 8.74% decrease in total pressure loss. At Mach number of 0.5 the total pressure loss decreases by 12.08%. At the Mach number of 0.67 near the stall incidence of +4°, the fluidic oscillator jet position at 7% of chord shows the most significant decrease in total pressure loss, which is 5.3%. When the pulsed jet is positioned at the separation starting point of the corner, the low-pressure region on the suction surface near the endwall decreases, while the high-pressure region on the pressure surface increases, enhancing the pressure expansion capability of the cascade. The pulsed jet can successfully decrease the vorticity in the passage, such as passage vortex, trailing shedding vortex and concentrated shedding vortex, and can resist the transverse secondary flow in the cascade passage, reducing the range of corner separation in the spanwise direction and total pressure loss.

Loss reduction in the compressor corner region via blade cooling

With Reynolds number (Re) decreasing to a critical value (approximately 2.0 × 105), the laminar-to-turbulent transition on the blade/endwall surface accelerates the loss generation in the corner region of compressors. Thus, how to effectively control the three-dimensional turbulent mixing loss at a low Re is of great importance. In this study, the mechanism of loss variation in a high-speed compressor cascade as Re decreases from 5.6 × 105 to 1.5 × 105 is investigated by large eddy simulations (LESs). Then, an isothermal cooled wall condition with a constant temperature of 0.8T01 (T01 is the inlet total temperature) is specified on the blade surface to explore the potential of loss reduction via wall cooling at Re=1.5 × 105. At Re=5.6 × 105, the boundary layers on the blade suction surface strongly interact with the endwall boundary layers and multiple passage vortex structures near the spanwise location of x/H = 0.3, inducing a large-scale concentrated shedding vortex (CSV) and determining the loss generation rate in the corner region. With Re decreasing to 1.5 × 105, the blade boundary layers grow much thicker due to the delayed transition. Detailed loss classification result indicates that the stronger interaction between the thickened blade boundary layers and secondary flow accelerates the local generation of turbulent fluctuations, which is mainly responsible for the performance deterioration in the corner region at a low Re. As the blade surface is cooled, the growth of blade boundary layers and the migration of secondary flow in the blade passage can be controlled simultaneously. Thus, the interaction between the blade boundary layers and secondary flow near the spanwise location of x/H = 0.3 is weakened, leading to a reduction in the overall total pressure loss by 13.3%. The results hope to guide in controlling the loss in the corner region of compressors operating at a low Re.

Experimental Study on Performance of Transonic Compressor Cascade with Microgroove Polyurethane Coatings

Due to the harsh operating environment of aero-engines, a surface structure that provides excellent aerodynamic performance is urgently required to save energy and reduce emissions. In this study, microgroove polyurethane coatings fabricated by chemical synthesis are investigated in terms of their effect on aerodynamic performance, which is a new attempt to investigate the impact on aerodynamic performance of compressor cascade at transonic speeds. This method reduces manufacturing and maintenance cost significantly compared with traditional laser machining. Wake measurements are conducted in the high-speed linear compressor cascade wind tunnel to evaluate the performance of cascade attached with different microgroove polyurethane coatings. Compared with the Blank case, the microgroove polyurethane coatings have the characteristic of reducing flow loss, with a maximum reducing rate of 5.87% in the area-averaged total pressure loss coefficient. The mechanism of flow loss control is discussed through analyzing the correlation between the total pressure distribution and turbulence intensity distribution. The results indicate that a large quantity of energy loss in the flow field due to turbulence dissipation and the reduction in viscous drag by microgroove polyurethane coatings relates to its effect on turbulence control. This paper demonstrates a great perspective on designing micro-nano surface structure for aero-engine applications.

Vortical characteristics of the corner separation flow in a high-speed compressor cascade

In aero-engine compressors, corner separations cause blockage, loss generation, and flow instabilities. The aim of the current study is to gain better knowledge of the vortical structures in corner separation flows and their resultant blockage and loss. Numerical simulations at different incidences have been performed for a high-speed linear compressor cascade. First, vortical structures have been visualized with the help of the Q-criterion and eigenvector method, and those associated with the corner separation have been focused on. It was found that the corner separation occurring on either the pressure surface or suction surface corner region was characterized by a large-scale streamwise vortex, named as the corner separation vortex in this paper. The pressure-side corner separation vortex is observed at a great negative incidence initiated from a focus point on the pressure surface, while the suction-side corner separation vortex observed at all concerned incidences originated from a focus on the endwall near the suction surface. Then, the mechanism by which corner separation vortices resulted in blockage and loss has been uncovered. The results showed that low-momentum fluids inside the core region of corner separation vortices led to the flow blockage and the viscous friction between them, and mainstream was responsible for the entropy generation. One major contributor of low-momentum fluids in the suction-side corner region was an off-wall reversed-flow region which was attributed to the breakdown of the suction-side corner separation vortex. Finally, multi-vortex structures in the leading region of the suction-side corner separation at two high positive incidences have been analyzed. It was found that the horseshoe vortex influenced the formation of the suction-side corner separation vortex, which induced a multi-vortex structure, and prevented the onset location of suction-side corner separation vortex from moving upstream with the increased incidence.

Loss reduction in a high-speed compressor cascade using an L-shaped endwall groove to generate a streamwise vortex

Reducing the endwall loss is essential for the development of high-efficiency aero-engine compressors. An advanced vortex generation technique, using an L-shaped groove in the endwall, has been applied to a high-speed linear compressor cascade to manipulate the complex endwall flow and hence reduce associated aerodynamic loss by means of an induced streamwise vortex. A geometry generation method for the L-shaped endwall groove has been developed for the purpose of facilitating the design of the L-shaped groove. Hereafter, the L-shaped endwall groove has been optimized through a surrogate-based optimization framework. The numerical results demonstrated that the optimal L-shaped groove could control the endwall flow effectively and achieve a significant loss reduction at all operating points that were considered. At the design point, the calculated total pressure loss coefficient was decreased by 17.9%, with the static pressure rise increasing by 5.53%. The flow underturning and loss away from the endwall were reduced, while the flow overturning and loss near the endwall increased. The analysis of the flow-control mechanism indicated that the groove separation vortex, i.e. the streamwise vortex generated by the L-shaped endwall groove, presented a barrier to the endwall cross flow, mitigating the interaction of the endwall fluid with the suction side of the blade.

Analysis of unsteady flow in compressor cascade based on modal DMD methods

The reduced-order model based on the modal decomposition method can capture the main characteristics of the compressor. In this part, dynamic mode decomposition (DMD) methods are used to analyze the flow in a high-speed compressor cascade. The methods can extract dynamic information of the flow field. Both the design and off-design conditions (-2°, 0°, 4°) are analyzed. The DMD methods can also obtain the global energy, time coefficient and spatial characteristics of the flow field. DMD mode energy of the characteristic frequency is the highest. The mode with the frequency doubling also has high energy at off-design points. The time coefficient obtained by DMD changes uniformly and eigenvalues are located on the unit circle, which indicates that there is no divergent or convergent mode.

Analysis of unsteady flow in compressor cascade based on modal POD methods

There are complex secondary flow structures in the compressor cascade passage, which seriously suppresses the performance of the compressor. The flow structures are highly unstable, especially for highly loaded compressor cascades. The reduced-order model based on the modal decomposition method can capture the main characteristics of the flow field and offer a reference for the application of unsteady flow control methods. In this study, the POD methods are used to analyze the flow in a high-speed compressor cascade. The methods can extract the mode that reflects the main energy of the flow field. Both the design and off-design conditions are analyzed. The decomposition methods can obtain the global energy, time coefficient and spatial characteristics of the flow field. The result showed that the flow field of the cascade has high steadiness and regularity. The zeroth POD mode under design conditions occupies over 99% of the energy of the flow field. The fluctuation velocity of off-design conditions is larger. The first five POD modes can contain over 90% of the energy of the fluctuation velocity field.

Study on the behavior of streamwise vortices formed between leading edge tubercles in a compressor cascade

A study has been carried out to investigate the formation mechanism and development of streamwise vortices induced by leading edge tubercles in a high speed compressor cascade. The preliminary assessment of the cascade performance in terms of the total pressure loss coefficient shows that the loss reduction is achieved at high incidence angles. A smaller wavelength leads to higher additional losses at the design point, but gives rise to a greater loss reduction at high incidence angles. The W4 cascade achieves the maximum loss reduction of 36.1% at i = 10°, as well as the stall angle improvement of 27.6%. The formation mechanism of streamwise vortices is elaborated on the basis of the streamwise vorticity equation, in which the streamwise turning terms may be responsible for the generation of streamwise vortices. Slices of streamwise vorticity at various streamwise locations, combined with vorticity strength distributions, have been presented to study the development of streamwise vortices. The counter-rotating vortices are divided into the crest-induced streamwise vortices (CSVs) and trough-induced streamwise vortices (TSVs). A streamwise vortex pair formed from part of the CSV sheets behind troughs, is gradually entrained by the TSV pair along the streamwise direction. In addition, the tubercles with a smaller wavelength result in higher streamwise vorticity strength, with which the streamwise vortices interact with the flow separation more sufficiently and delay the separation to a greater extent.

The Performance of the Self-Supplying Vortex Generator Jets on a High-Speed Compressor Cascade

Abstract The self-supplying vortex generator jet (SVGJ) is introduced to the separation control of a high-speed compressor cascade with an inlet Mach number of 0.67. A parametric investigation concerning the jet location and the injection angle is conducted numerically. The effect mechanisms of SVGJ on the cascade performance are analyzed in detail. The potentials of SVGJ at off-design points are also discussed. The vortex generated by the jet utilizing the pressure difference between the pressure surface and suction surface could enhance the fluid exchange between the boundary layer and the mainstream, thus weakening the flow separation and reducing the flow loss effectively. With a hole of diameter 0.5 mm and a jet-to-inflow mass flow ratio less than 0.3‰, the maximum loss reductions are 5.2 % at the design point and 8.0 % at the off-design point, validating the high efficiency of SVGJ. In this work, the optimum performance is obtained by the SVGJ located at xj /B=40 %, h/H=15 % with a skew angle of β=60°.

Numerical loss analysis in a compressor cascade with leading edge tubercles

A numerical analysis of loss has been carried out to explore the loss mechanism of leading edge tubercles in a high speed compressor cascade. Taking the lead from flippers of the humpback whale, tubercles are passive structures of a blade for flow control. Evaluation of the overall performance in terms of entropy increase shows that the loss reduction is achieved both at high negative and high positive incidence angles, while a rise in the loss is obtained near the design point. And a smaller wave number as well as a smaller amplitude results in lower additional losses at the design point. Spanwise and streamwise distributions of pitchwise-averaged entropy increase combined with flow details have been presented to survey the loss development and, subsequently, to interpret the loss mechanism. The tubercle geometry results in the deflection flow and the consequent spanwise pressure gradient. This pressure gradient induces formation of counter-rotating streamwise vortices, transports away the low-momentum fluid near wall from crests towards troughs and leads to local high loss regions behind troughs as well as loss reduction behind the crests in comparison to the baseline. The interaction between these vortices and flow separation by momentum transfer leads to separation delay and the consequent loss reduction at the outlet.

A Parametric Investigation of Wake Loss Reduction by using Serrated Trailing Edge in a High-speed Compressor Cascade

Abstract The influence of the serration with different widths, angles and spanwise locations on the wake of a high-speed compressor cascade is parametrically investigated by numerical simulation. The effect mechanisms of the serration on the cascade performance are discussed in detail and the potential at the off-design points is also validated. The results show that the momentum inputted by the serration jet and the fluid exchange induced by the dominant jet vortex could reenergize the low energy corner region and weaken the trailing edge shedding vortex. With the upstream shift of the serration valley, the loss reduction and pressure rise are enhanced due to the strengthened jet and counter-rotating vortex pair, whereas the flow turning is reduced by the decreased camber angle and the enhanced jet deflection. The serration at the blade height corresponding to the core of the low energy fluid could perform better for the loss reduction, whereas the serration close to the endwall could obtain higher pressure rise. For the optimal serration located at 12.5 % blade height with a width of 5 mm and an angle of 45 °, the loss is reduced by 8.9 % and the pressure rise is increased by 2.2 %, with the flow turning slightly reduced by 0.93 °

Secondary flow control using endwall jet fence in a high-speed compressor cascade

This paper proposes a secondary flow control concept using Endwall jet fence (EJF). A parametric investigation concerning the variations of the jet location along the axial and pitch-wise direction as well as the skew angle is conducted numerically to validate the potential of EJF in a high-speed compressor cascade with an inlet Mach number of 0.67. And then the interaction mechanisms between the EJF and the endwall secondary flow are discussed in detail. The results show that the EJF could reduce the corner separation and losses significantly by inputting transverse momentum component, inducing a concentrated jet vortex to block the pitch-wise migration of the passage vortex as well as enhancing the energy exchange between the endwall boundary layer and the mainstream. The jet location and the skew angle are important for the influence of EJF on the cascade performance. In this work, a maximum total pressure loss reduction of 11.6 % is obtained by the EJF located at 30 % of the axial chord and 10 % of the pitch with a skew angle of β = 40°, whereas the jet-to-inflow mass flow ratio is only about 0.4 %, validating the high efficiency of this flow control concept. For the off-design points, the EJF also shows appreciable potential on the endwall secondary flow control and loss reduction.

Control of Corner Separation with Plasma Actuation in a High-Speed Compressor Cascade

The performances of modern highly loaded compressors are limited by the corner separations. Plasma actuation is a typical active flow control methodology, which has been proven to be capable of controlling the corner separations in low-speed compressor cascades. The main purpose of this paper is to uncover the flow control law and the mechanism of high-speed compressor cascade corner separation control with plasma actuations. The control effects of the suction surface as well as the endwall plasma actuations in suppressing the high-speed compressor cascade flow separations are investigated with numerical methods. The main flow structures within the high-speed compressor cascade corner separation and the development of the corresponding flow loss are investigated firstly. Next, the performances of plasma actuations in suppressing the high-speed compressor cascade corner separation are studied. At last, the mechanisms behind the control effects of the suction surface and the endwall plasma actuations are discussed. Both the suction surface and the endwall plasma actuations can improve the high-speed compressor cascade static pressure rise coefficient, while reducing the corresponding total pressure loss and blockage coefficients. The suction surface plasma actuation can suppress not only the high-speed compressor cascade corner separation vortex but also the airfoil separation, so, compared to the endwall plasma actuation, the suction surface plasma actuation is more efficient in reducing the total pressure loss of the high-speed compressor cascade. However, through suppressing the development of the passage vortex, the endwall plasma actuation is more efficient in reducing the flow blockage and improving the static pressure rise of the high-speed compressor cascade.

Separation control with endwall corner jet in a high-speed compressor cascade

Separation control with endwall corner jet in a high-speed compressor cascade

Impact of Weld Joint on Compressor Blade Performance

This contribution focused on the analysis of weld-joint effects resulting from the assembly of the blades and disk for the compressor blades. Experiments at the von Kármán Institute and computational fluid dynamics simulations at Cenaero were conducted to study the flowfield in a low-pressure high-speed compressor cascade with two weld-joint configurations. Experiments were performed to measure the total pressure loss coefficient for various inlet flow conditions. Reynolds-averaged Navier–Stokes simulations using the turbulence model were carried out. Comparison between the Reynolds-averaged Navier–Stokes solutions and the experiments were made, and the accuracy of the results was investigated. The impact of the weld-joint height and the upstream boundary layer on the losses was studied. In addition, the effect of the weld-joint heights on the growth of the boundary layer was analyzed and presented.

Experimental investigation on a high subsonic compressor cascade flow

With the aim of deepening the understanding of high-speed compressor cascade flow, this paper reports an experimental study on NACA-65 K48 compressor cascade with high subsonic inlet flow. With the increase of passage pressurizing ability, endwall boundary layer behavior is deteriorated, and the transition zone is extended from suction surface to the endwall as the adverse pressure gradient increases. Cross flow from endwall to midspan, mixing of corner boundary layer and the main stream, and reversal flow on the suction surface are caused by corner separation vortex structures. Passage vortex is the main corner separation vortex. During its movement downstream, the size grows bigger while the rotating direction changes, forming a limiting circle. With higher incidence, corner separation is further deteriorated, leading to higher flow loss. Meanwhile, corner separation structure, flow mixing characteristics and flow loss distribution vary a lot with the change of incidence. Compared with low aspect-ratio model, corner separation of high aspect-ratio model moves away from the endwall and is more sufficiently developed downstream the cascade. Results obtained present details of high-speed compressor cascade flow, which is rare in the relating research fields and is beneficial to mechanism analysis, aerodynamic optimization and flow control design.

Evaluation of RANS turbulence models in simulating the corner separation of a high-speed compressor cascade

A typical high-speed compressor cascade flow is simulated using an eddy viscosity model and a Reynolds stress model. Based on comparisons with experimental results, the simulation accuracies of the two turbulence models are evaluated, and some recommendations are provided for further research. Inside the cascade flow passage, the simulation accuracies of both turbulence models decrease because of the corner effect. Downstream of the blade, the wake mixing influence, and thus the flow loss, is under-predicted by both turbulence models. Compared with the linear eddy viscosity model, the modification of the simulation accuracy with regard to endwall flow prediction using the Reynolds stress model is limited, which only occurs at a highly positive incidence angle. The corner separation strength and wake mixing influence predicted by the eddy viscosity model are surpassed by those of the Reynolds stress model. At a negative incidence angle, the simulation result of the Reynolds stress model is proved to be physically unreasonable according to the flow loss prediction.