فهرست مطالب z. x. wang

To enhance the aerodynamic performance of an ultra-compact S-shaped convergent-divergent nozzle and mitigate flow separation, numerical simulations were conducted using FLUENT software. The study employed the k-ω shear stress transport turbulent model to investigate a flow control method involving blowing. Detailed analysis was performed on the impact of blowing position, angle, and pressure ratio on controlling flow separation. The findings indicate that as the blowing position moves backward, the flow separation area diminishes. Additionally, downstream flow separation ceases at smaller blowing angles within the separation zone. However, excessively large blowing angles tend to create an “aerodynamic wall,” causing significant upstream flow loss and nozzle performance degradation. Enhancing the blowing pressure ratio, given proper mixing with low-energy fluid and no interference with the main flow, can improve the nozzle's aerodynamic performance. Under the optimal blowing scheme, the total pressure recovery coefficient and thrust coefficient are increased by approximately 0.52% and 3.75%, respectively, when compared with those of the reference nozzle.

Serpentine nozzle can effectively suppress the infrared radiation signatures of the aero-engine exhaust system. However, it experiences the remarkable fluid-structure interaction (FSI) process at the work condition. In this paper, the deformation behavior of the serpentine nozzle and its flow characteristic were investigated numerically. Then, the influences of the wall thickness and the geometric configuration on the FSI effect were also explored. The results show that, the mechanism of the fluid-structure interaction is formed through the data transfer of the force and the displacement at the FSI interface. Under the FSI effect, there occur the ballon-like swellings at the second S passage, and the linear section bends upward along the Y direction. They induce the special flow features including the flow separation, the shock wave and the plume vector angle. As the value of the wall thickness increases from 3mm to 6mm, the maximum of the deformation displacement of the serpentine nozzle decreases 68.5mm. As compared to the uncoupled state, the variation of the axial thrust decreases from 2.70% to 0.70% at the coupled state. The circular-to-rectangular profile and the S-shaped passage enlarge the deformation behavior of the nozzle structure. The value of the axial thrust of the serpentine nozzle with 5mm wall thickness for the coupled state is lower 1.92% than these for the uncoupled state.

To understand the jet flow characteristics of turbofan separate exhaust system, a parametric design method based on the initial Class Shape Transformation function was developed. HBPR and UHBPR turbofan separate exhaust systems were designed. Furthermore, the jet flow characteristics of the HBPR turbofan exhaust system under take-off condition with zero angle of attack were studied based on numerical simulation. The jet flow characteristics of the HBPR and UHBPR turbofan exhaust system under take-off condition with high angle of attack were also simulated. The effects of angle of attack and bypass ratio on the jet flow characteristics were investigated and the related flow mechanisms were analyzed. Results show that the axisymmetric plumes of the HBPR turbofan exhaust system are distributed around the engine axis under take-off condition with zero angle of attack. With the plug wake as the center, the core flow, the fan/core shear layer, the fan flow, the fan/free stream shear layer and the free stream are wrapped around the plug wake from inside out. Vortexes appear in the lee area at the back of the cowl and jet flow under take-off condition with high angle of attack. These vortexes cause cross sectional secondary flow and expose the high-velocity core flow to the low-velocity free stream. The contact area and velocity gradient in the mixing region among the free stream, fan flow and core flow increase. Therefore, the mixture among jet flow and free stream strengthens. So the high-velocity region, the high-vorticity region, and the high turbulence kinetic energy region shorten by 55.1%, 47.7% and 50.9% respectively. The vorticity values and turbulence kinetic energy level peak on the upper side of the exhaust plumes increase by about 30% and 87% respectively. Relative to these parameters from the HBPR turbofan exhaust system, the jet velocity peak value of UHBPR turbofan decreases by 5.5% under take-off condition with high angle of attack. The vorticity values and turbulence kinetic energy level reduce due to decreased velocity gradient in shear layers downstream of the nozzle exit plane. The turbulence kinetic energy level peak on the upper side of the exhaust plumes decreases by 29.3%. The reasons are that the contact area between high-velocity core flow and the free stream decreases due to thicker fan flow and the velocity gradient in the core flow and free stream mixing region decreases because of the lower core flow velocity.

To assess unsteady vortex interaction between rim seal purge flow and upstream stator, numerical ‎investigations were conducted under different purge flow rates. The vortex distributions for the stator and ‎cavity were investigated and the interaction processes near the cavity exit, in particular the vorticity ‎change resulting from the ingress and egress, were analyzed. Results show the intensity of hub passage ‎vortex (HPV) and hub trailing shedding vortex (HTSV) at stator exit is decreased as a consequence of ‎enhancing blockage effects caused by the egress flow. However, when the purge flow rate increases, ‎from stator exit to downstream of cavity exit, the reduction in the intensity of two vortices is weakened as ‎the extrusion of egress flow thins their vortex tubes. The vortex inside the cavity is generated as the ‎combined effect of relative rotation of cavity walls and non-uniform circumferential pressure mainly ‎imposed by upstream stator. The ingress leads the positive axial vorticity near the stator hub to ingest into ‎the cavity and eject into the main passage due to the blockage of purge flow. Furthermore, the interaction ‎between the ingress of the mainstream and purge flow produces local negative axial vorticity. The egress ‎flow carries negative axial vorticity mainly originated from the rotational cavity wall, and enters into the ‎main flow passage near the rotating hub, then locations of HPV and HTSV move to the mid-span slightly ‎with the extrusion of egress flow‎.

Lighter weight, simpler structure and throat area controllable are the developing trends of aircraft engine exhaust system. To meet these challenges, a new concept of hybrid throat control (TC) nozzle was proposed to improve the control efficiency of throat area (η) by using a rotary valve with secondary injection. The flow mechanism of the hybrid TC nozzle and the effect of aerodynamic and geometric parameters on nozzle performance were investigated numerically. Then the approximate model characterizing the hybrid TC nozzle was established with design of experiment and response surface methodology. The approximate model was used to analysis the coupling effect between parameters and optimized the parameter combination. The results show that the flow area of the nozzle can be restricted effectively by the rotary valve and the secondary flow, and η is bigger than 5.24. Nozzle pressure ratio and secondary pressure ratio are the dominant factors for the nozzle throat area control performance. The optimization of the parameter combination was carried out with penalty function approach, with ratio of throat area control being 30 percent and corrected mass flow ratio of secondary flow being 5 percent The maximize error of the optimization result is 4.13 percent and it verifies the validity and feasibility of the approximate model.

In order to understand the coupled effect of the nacelle and exhaust system and to improve their overall performance, we studied the aerodynamic performance and the flow characteristics of the high bypass ratio turbofan nacelle and exhaust system by numerical simulation. The geometric parameters of a nacelle and exhaust system (e.g., the contraction ratio of the cowl afterbody and the fan nozzle exit angle) were investigated to evaluate their influence on the overall performance of the nacelle and exhaust system. The related flow mechanism was explored as well. The results show that the flow field of the nacelle and exhaust system under the mid-cruise condition exhibits characteristics of transonic flow. A stagnation zone exits at the nacelle lip and there is a velocity peak at the nacelle forebody. There exist a number of complex flow phenomena (such as shockwave, expansion wave, shear flow and shock wave-boundary layer interaction) in the downstream of the fan nozzle exit plane. The magnitude of the fan nozzle thrust or the intake ram drag is much higher than that of the additional drag, the nacelle drag or the core nozzle thrust. And for the nacelle drag, the friction drag of the cowl is in the same order of magnitude as the pressure drag of the cowl, the core cowl and the plug. But it is much larger than the friction drag of the core cowl and the plug. The effective thrust increases by 4.7% as the contraction ratio of the cowl afterbody increases; and it increases by 2.4% as the fan nozzle exit angle increases. The expansion degree of the fanjet flow, the shock wave strength and location, and the existence of the flow separation or second shock wave are influenced by the contraction ratio of the cowl afterbody and the fan nozzle exit angle. These phenomena have effects on the pressure distribution of the core cowl and the surrounding fanjet flow velocity, and hence they further affect the nacelle drag. The increase in the fan nozzle exit angle can noticeably reduce the thrust of the fan nozzle.