Volume 42, Issue 3

Numerical simulations of a three-dimensional asymmetric nozzle with lateral expansion were carried out, focusing on the characteristics of the three-dimensional flow field in the over-expanded state and the influence of the sidewall surface profiles on the over-expanded flow field. The results show that the sidewall profiles cause the separation zones inside the nozzle to exhibit different shapes. The center separation bubble and the corner zone separation bubble are independent of each other in the lower wall surface confined shock wave separation mode. In the upper wall-constrained shock wave separation mode, the center separation bubble, the corner zone separation bubble, and the sidewall surface separation bubble are interconnected. Under the same drop pressure ratio, changes in the sidewall profile affect the shape and size of the separation zones, and the convex wall can slow the occurrence of separation to a certain extent.

In this study, parametric optimization is employed to identify optimal output parameters for physical components of HOMER nozzle type fluidic thrust vectoring (FTV) system. Optimization study is conducted on seven output parameters: moment around upper Coanda surface, M A ; performance parameter, PP; thrust vectoring angle, θ T ; total thrust, T; thrust vectoring efficiency, η ; axial thrust, T x ; transverse thrust, T y . 440 case studies in design-of-experiment stage are employed, utilizing a range of meta-models. Genetic-Aggregation meta-model exhibited highest coefficient of determination (R 2 ) among tested meta-models. In response surface optimization stage, six objectives are enforced through use of Multi-Objective-Genetic-Algorithm technique. Subsequently, three optimization candidate points are identified for filtering of output parameters deemed to be of particular importance. Among them, candidate point with geometry-4 and parameter-1 = 39.23 m/s, exhibited the highest values for PP, θ T , η, and T y , with 4.78 %, 4.48 %, 7.97 %, and 7.94 % higher values, respectively, compared to the baseline geometry.

In order to address the issue of excessive vibration and critical speed approaching the operating speed of the aero-engine fan rotor, the integral squeeze film damper (ISFD) was investigated for its impact on the fan rotor tester of turbofan. The static characteristic and dynamic characteristics of ISFD and fan rotor were analyzed through simulation calculation. The results show that ISFD can effectively suppress the vibration, improve the stress distribution within the rotor, and adjust the critical speed of the rotor. Experimental tests were conducted on the fan rotor tester of turbofan, revealing significant reductions in vibration: At 8,000 r/min, the peak-to-peak displacement was reduced by 25.35 %, while the vibration of 1X frequency and 2X frequency was suppressed by 84.48 % and 33.46 %, respectively. It also successfully adjusted critical speed within this system, reducing the critical vibration by 58.81 % with a decrease in critical speed from 7,397 r/min to 3,265 r/min.

To achieve a highly efficient conical shaft design in aviation gas turbine engine’s rotor system, firstly, a 4-DOF cantilevered rotor model is established to investigate the effects of the lateral, angular and cross stiffness of the rotor on the resonance speed of the pitching mode. The results show that increasing the lateral and angular stiffness, while reducing the cross stiffness, is beneficial for raising the critical speed of the pitching mode. Subsequently, using finite element method, the relationships between the cone angle, dimensionless radius, dimensionless span, and the stiffnesses of conical shaft are calculated. It is observed that, under various structural design constraints, the conical shaft with maximum stiffness exhibits different cone angles. Finally, experiments are conducted with flexible rotors featuring conical shafts of varying cone angles to validate the research conclusions. This paper provides guidance for the configuration design of rotors in aero-engine.

Dislocation is defined as the inability of shrouded blades to return to their normal position after deformation caused by an external force. Numerous methods exist for designing shrouds of turbine engines, targeting structural strength, sealing performance, contact interface wear, or damping. However, these designs often fail when dislocation occurs, which is a common issue in engineering. Limited methods focus on preventing shroud dislocation. In this article, we classify the dislocation of zig-zag shrouds into two patterns. A method for calculating the equivalent stiffness of the contact interface is provided to determine the clearance of non-working surfaces. Through a series of static analysis, we introduce a method to adjust the shroud shape and assess its capability in preventing dislocation. All methods are tested on a low-pressure turbine blade.

The purpose of this paper is to suppress the tip leakage vortex and cavitation of fuel pump by using the casing slot. The flow fields of the blades with different casing slots were studied by numerical simulation under the conditions of no cavitation and cavitation respectively. The results show that the loss with leakage vortex cavitation increases by 37 % compared with the condition without cavitation. Under cavitation conditions, the optimal scheme is c6 scheme, and the loss is reduced by 9.3 %. The tip leakage vortex and vortex cavitation can be effectively reduced by casing slot. Under the condition of no cavitation, the tip leakage flow loss can be effectively reduced by the casing slot in a wide range of axial positions. After the casing slot is applied, the shape of the blade tip leakage vortex is significantly changed, the size of the leakage vortex is significantly reduced, and the loss is reduced. Under cavitation conditions, the casing slot can also effectively improve the tip flow field and inhibit leakage vortex cavitation. In a wide range of axial positions, the casing slot can effectively inhibit leakage vortex cavitation.

This paper numerically studies the heat transfer and flow loss of array-jets impingement cooling with initial crossflow under three aerodynamic parameters (jet Reynolds number (Re j ), crossflow-to-jet mass flow ratio ( CMFR ) and crossflow-to-jet temperature ratio ( TR )) and four typical geometrical parameters (jet-to-target spacing( H ), jet-to-jet streamwise spacing ( P x ), jet-to-jet spanwise spacing ( P y ), and jet angle ( θ )). From the results, as Re j increases, the heat transfer is improved, and the flow loss is declined. The TR has obvious influence on heat transfer and flow loss, and the increased TR leads to a decrease in heat transfer, but the effect of TR on friction loss is closely related to specific geometrical parameters. The effect of CMFR on heat transfer and flow loss is similar to that of TR, both of which are related to specific geometric parameters. Through detail analysis of multi-parameters effect, this paper finally establishes the improved correlation formulas for area-averaged Nusselt number and friction loss coefficient under the “three-temperature” problem, which further lays a solid theoretical foundation for array-jets impingement cooling design.

This study investigates the impact of shockwaves induced by various combustor wall geometries – wedge, wavy wall, circular bumps, and triangular bumps – on scramjet combustor performance. Using CFD simulations in ANSYS Fluent 23.1 with RANS equations and the SST turbulence model, key parameters such as velocity, static pressure, temperature, turbulence intensity, and combustion efficiency are analyzed. Circular bumps exhibit superior performance, achieving higher turbulence intensity (42 %), enhanced fuel-air mixing, and minimized static pressure losses (8 %), optimizing supersonic flow dynamics. This geometry ensures better combustion efficiency, reducing unburned fuel and maximizing heat release. While wedge and triangular geometries improve mixing and flow stability, they are less effective than circular bumps. The wavy wall structure provides a balanced performance. The findings highlight the potential of circular bumps in advancing scramjet combustor designs, offering valuable insights for hypersonic propulsion system enhancements.

In existing calculation studies of gas turbine part-load performance, the multi-stage power turbine (PT) is usually considered as an integral whole, and the variation rule of power and efficiency of the PT’s front/rear stage with load cannot be obtained. Therefore, the 6-stage PT of LM2500+ gas turbine is divided into 2 sub-turbines to be investigated by establishing the gas turbine part-load performance model based on the stage by stages method of the power turbine in this paper. The model is used to calculate the part-load performance of the LM2500+ under the fuel flow strategy (FF) and the first sub-turbine variable area nozzle (VAN) strategy. The results show that the relative errors of the mass flow rate, pressure ratio and thermal efficiency calculated by the model compared to the experiment data are all within 3.5 %.Under the FF strategy, the degree of mismatch in PT front/rear stage gradually increases as the load decreases from 100 % to 0.0 %, with the share of the first sub-turbine power increases from 22 % to 130 %, and the second sub-turbine efficiency decreases from 0.876 to −0.126. The adoption of the VAN strategy further exacerbates the degree of mismatch between the PT front/rear stages.

The reduction in the potential core length and the enhancement of mixing by varying the lip thickness with co-flowing jets through a rectangular nozzle have been investigated using both computational and experimental methods. The lip thickness of a nozzle is defined as the thickness of the primary nozzle wall that separates the primary and secondary co-flowing jets at the nozzle exit. A baseline rectangular nozzle with a hydraulic diameter of 6 mm and three co-flow nozzles with annular co-flows with lip thicknesses (LT) of 25 %, 50 %, and 75 % of the baseline hydraulic diameter were designed and studied for a subsonic jet Mach number of 0.8.The jet centerline pressure decay, radial total pressure decay, and jet mixing behavior were analyzed. The results show that increasing the lip thickness enhances the mixing characteristics of the primary and secondary flows. The reduction in the potential core length was 39.39 % for the 75 % LT configuration compared with that of the plain convergent rectangular nozzle, and 7.27 % and 19.24 % for the 25 % and 50 % LT configurations, respectively.

Emission characteristic of a single annular gas turbine combustor has been studied at engine part-load conditions from idle to max power setting. Under fuel-rich condition. The temperature field and pollutant formation at different power settings are studied in detail. The equivalence ratio varies from 0.35 corresponding idle power setting to maximum 1.1 at max power setting at take-off condition. Concentration of major pollutant species at combustor exit such as NOx, CO, CO 2 , soot, unburned hydrocarbons and water vapor are analyzed for varying inlet conditions and primary zone equivalence ratios at different power settings using model in ANSYS CFX. The pollutant concentrations are found to be increased at combustor exit with increase in power setting as equivalence ratio and flame temperature increase. The numerical predictions at low regime of power setting are validated against gas sampling data obtained from actual combustor testing in aerothermal test facility. The numerical analysis based on satisfactory validation with limited experimental data is very reasonable and provides input for prediction of pollutant emissions of the intended derivative engine during the landing-takeoff (LTO) cycle prior to engine test.

Two-stage floating ring seal is a non-contact circumferential seal with low leakage and high reliability. This paper elaborates on the variation mechanism and calculation method of the working clearance and working conditions of sealing chamber, reveals the corresponding relationship between operating conditions of aircraft engine and leakage characteristics of two-stage floating ring seal, obtains an effective method for analyzing the leakage characteristics of two-stage floating ring seals. The pressure of sealing chamber can be determined by continuously assuming the pressure of sealing chamber and substituting it to continuity equation. Carry out leakage characteristic experiments under specific operating conditions, the minimum deviation between the value of calculation and the experimental results is −11.47 %, the maximum deviation is 7.34 %, and the average of the absolute value of deviation is 5.01 %. The optimization design scheme for the working clearance and the intermediate air flow channel have been proposed, which can provide important support for improving the reliability and sealing performance of the two-stage floating ring seal.

A boundary layer analysis is presented for hypersonic flow at an axisymmetric stagnation point region of a blunt body under isothermal and adiabatic boundary conditions. Consideration is given to variable properties of air. It has been shown that surface drag and heat transfer rates may be controlled by applying magnetic field and vectored surface mass transfer. The range of Mach numbers considered is 1–10. As the magnetic field strength M increases, friction factor and heat transfer rate (in the case of isothermal surface) or surface temperature (in the case of adiabatic surface) increase.Friction factor and surface temperature (in the case of adiabatic surface) can be reduced by applying vectored surface mass transfer.

Numerical investigation of turbulent lifted H 2 /N 2 jet flame in vitiated hot coflow is presented. Turbulent combustion modeling is performed by using flamelet generated manifold approach and presumed probability density function for turbulence-chemistry interaction. Sensitivity analysis of scalar dissipation modeling for mixture fraction and progress variable suggests that the combustion temperature is more sensitive to mixture fraction variance, whereas the flame lift-off height is more sensitive to progress variable variance. The present simulations provide overall good agreements with experimental measurements for mixture fraction, temperature, and chemical species, which gives further information regarding the global flame structure and reaction characteristics. Investigation on mixture fraction conditional scatter plot reveals that the evolution of lifted flame branch and flame dynamics are well captured. Two combustion modes are identified in flame zone. It indicates that non-premixed combustion dominates in core flame region, however low probability of premixed combustion is observed occurring in central fuel-rich zone.

In this study, the Kriging model is employed alongside the Efficient Global Optimization (EGO) method to optimize the sealing performance of a labyrinth seal structure. Two optimization strategies are explored: one involving the consistent variation of design parameters and the other, independent variation. A comparative evaluation of the two approaches demonstrates their effectiveness, followed by an in-depth analysis of the leakage characteristics under varying operating conditions, both before and after optimization. Additionally, a sensitivity analysis is conducted for the independent optimization scheme to examine the influence of design variables on the objective function. The results indicate that both optimization schemes significantly improve the sealing performance, with the equivalent flow rate reduced by 14.5 % and 18.47 %, respectively, compared to the baseline design. The optimized models maintain a sealing advantage under variable operating conditions. Furthermore, the sensitivity analysis reveals key design parameters influencing leakage performance, providing valuable insights for the structural optimization of labyrinth seals.

This study aims to evaluate the film-cooling performance of trench film cooling with a modified cooling hole injection angle configuration. Five different film hole injection angle configurations viz., forward injection 30°, forward injection 60°, perpendicular injection 90°, backward injection 60°, and backward injection 30° were tested for their performance through Computational fluid dynamic simulation. All the configurations were tested on a trench with a fixed length of 2 times and depth of 0.75 times the film hole diameter. The investigation was conducted at three blowing ratios of 0.4, 1.2, and 2.0. The experimental results of trench film cooling effectiveness with a backward injection of 30° show a close agreement with the computational data. The backward injection 30° film hole configuration has improved the spatially averaged film cooling effectiveness on the trench test surface by 29.41 % at a blowing ratio of 0.4, 30.65 % at a blowing ratio of 1.2, and 47.18 % at a blowing ratio of 2.0.

To suppress corner separation, numerical simulation methods were used to investigate the effects of two passive flow control technology, non-axisymmetric endwall profiling (NAEP) and vortex generator (VG), on the aerodynamic performance of a diffusion cascade under design conditions. The results indicate that under the D1 circumferential installation position, the optimal combination of the vortex generator stagger of 15° and the NAEP of the blade suction surface concave structure in the passage is optimal. Under design conditions, the VG has a significant effect on suppressing the development of laminar separation bubbles in the middle of the chord, while NAEP has a worse effect. However, the NAEP scheme has a significant effect on the corner separation and reversed flow in the corner region, reducing the Cpt by 28.433 %. The combination scheme combines the advantages of both control technologies. Under design conditions, the high dissipation region caused by laminar separation bubble migration on the suction surface of the blade is reduced, the separation in the corner region is further suppressed, the reversed region in the corner region is reduced, the diffusion capacity of the cascade is improved, and the Cpt of the combination scheme is reduced by 29.045 %.

The experimental findings on the properties of an suddenly expanded flow from convergent nozzle is presented in this article. For suddenly expanded subsonic, sonic and underexpanded flows, the effect of passive control in the form of an annular rib on the base pressure distribution along the enlarged duct is examined. The range of the nozzle pressure ratio (NPR) was 1–7. The enlarged duct’s length-to-diameter ratio was adjusted from the base to 0.5D to 6D. The rib’s aspect ratio was adjusted between 0.45 and 1.25. The experimental findings demonstrate that, in the absence of ribs, the base pressure is highly impacted by the flow Mach number, the degree of expansion at the nozzle exit, and the length-to-diameter ratio of the enlarged duct. The base pressure of a sonic jet with NPR 4 rises as the rib aspect ratio does as well. The current study clearly shows that, in a suddenly expanded axi-symmetric duct, the base pressure can be managed passively by using a rib with the proper aspect ratio. This prevents wall pressure oscillations from rising to an unacceptably high level in subsonic and appropriately and underexpanded sonic flow.

Understanding how preheated multi-annular swirling air jets mix within confinements of different expansion ratios is crucial. This knowledge is essential for applications like gas turbine combustors, where efficient mixing and maintaining flame stability are essential for optimal performance. The central recirculation zone (CRZ) is crucial in improving mixing in swirling confined flows. This investigation focuses on the computational analysis of the mixing and flow characteristics of three swirling air jets, considering the effects of varying expansion ratios and preheating of the jets within different confinements. The study utilized the CFD software ANSYS FLUENT to investigate the impact of jet preheating on the recirculation zone and flow behavior within confinement. The Realizable k-ε turbulence model was employed, and the CFD code was validated by comparing its results with experimental data available in the literature. In the present investigation, it was found that the preheating of jets influences the size of the recirculation zone at a specific swirl combination. The findings offer valuable insights into the mixing behavior of preheated jets in both non-expanded and expanded confinements.