Low-observable serpentine nozzles are essential for the exhaust systems of new-generation stealth aircraft. Considering the complex spatial constraints and compact layout of serpentine nozzles, dual-stream serpentine nozzles with varying geometric characteristics were designed. The influence of three key geometric parameters, namely the nozzle exit aspect ratio (W_e/H_e), length-to-diameter ratio (L/D_in), and obscuration ratio (△Y), and their interactions with the structural deformation of the serpentine nozzle were investigated based on a sequentially coupled two-way Fluid-Structure Interaction (FSI) algorithm. Response surface models for structural deformation were established using the Box-Behnken Design (BBD) experimental design method and Response Surface Methodology (RSM), and the sensitivity of key geometric parameters to the deformation of critical nozzle locations was analyzed. The results indicate that each geometric parameter has varying degrees of influence on the deformation of different nozzle sections, and the coupling effects between parameters are significant. W_e/H_e primarily affects the deformation of the top exit wall, △Y mainly influences the deformation of the bottom exit wall and the downstream wall of the first bend, while L/D_in has a certain degree of influence on the deformation of all sections. With the objective of minimizing structural deformation while maintaining good aerodynamic performance, the serpentine nozzle was optimized using RSM, resulting in a 56.9% reduction in the total structural deformation.

To investigate the influence and generation mechanism of design parameters on the infrared radiation characteristics of serpentine nozzles under complex flow conditions generated in the mixing chamber of aviation engines, this study employs numerical simulation methods to calculate the infrared radiation characteristics of serpentine nozzles with different aspect ratios and length-diameter ratios. All serpentine nozzle models meet the constraint of completely shielding the high-temperature components. Additionally, the study analyzes the coupling mechanism between complex non-uniform flow and infrared radiation characteristics with different design parameters. The research findings indicate that the longitudinal vortex generated by the lobed mixer in the mixing chamber entrains high-temperature gas impacting the nozzle wall, leading to a large area of hot spots appearing on the upper and lower walls of the serpentine nozzle. On the vertical detection plane, the total infrared characteristic level increases by 1.22 times compared to that on the horizontal detection plane. As the aspect ratio increases, the hot spot area enlarges due to the intensified longitudinal vortex attachment, while the jet length decreases because of the enhanced atmospheric mixing. Consequently, the total infrared characteristic level on the vertical detection plane remains nearly unchanged, whereas that on the horizontal detection plane decreases by 30.5% due to the minor influence of hot spots. When the length-diameter ratio is small, flow separation occurs inside the serpentine nozzle due to the large curvature of the wall. The intensity of the hot spots on the wall surface significantly increases, causing the total infrared characteristic level on the horizontal and vertical detection planes to increase by 23.5% and 38.6%, respectively, compared to the cases with-out flow separation. With further increase in the length-diameter ratio, the hot spots on the wall expand due to the full mixing of the core flow and the bypass flow, resulting in 15.0% and 29.4% increase in total infrared characteristic level on the horizontal and vertical detection planes, respectively.

The internal flow of the serpentine nozzle is complex, with numerous parameters affecting its aerodynamic performance. Traditional component-based zero-dimensional engine models cannot accurately assess the aerodynamic performance impact of the serpentine nozzle on the overall engine. In this paper, a high-fidelity performance prediction model for serpentine nozzles is established using the back propagation (BP) neural network, and is coupled with a zero-dimensional turbofan engine model. This integrated approach is employed to investigate the influence of the baseline serpentine nozzle on engine speed, altitude characteristics, and component behaviors, as well as the differences in nozzle and engine performance with various geometric parameters. The results indicate that compared to axisymmetric nozzles, the engine equipped with the serpentine nozzle experiences a decline in performance. Specifically, at sea-level static conditions, the engine's thrust decreases by 4.50%, while the fuel consumption increases by 4.75%; the fan bypass ratio decreases by a maximum of 0.33% at sea level, accompanied by a reduction in surge margin. Conversely, at an altitude of 12 km, the fan bypass ratio increases by a maximum of 0.28%, and the surge margin increases. These variation trends in fan operating characteristics can be attributed to the differences in throttling effects of the serpentine nozzle on the core and bypass of mixing chamber at different altitudes. Additionally, increasing the length-to-diameter ratio of the serpentine nozzle from 2.2 to 3.0 enhances its performance, resulting in an 8.0% increase in thrust coefficient and a 4.8% increase in discharge coefficient. The multi-dimensional coupling model between serpentine nozzle and engine established in this study can effectively evaluate the changes in engine performance and component characteristics following the installation of serpentine nozzles of varying geometric parameters.

A study was conducted on the aerodynamic interference effects of Contra-Rotating Propfan (CRP) and propfan engine inlet. Unsteady numerical calculation was carried out by the sliding grid method, and the coupled aerodynamic effects of the 8X8 CRP and inlet were compared and analyzed under two typical operating conditions: the high-altitude cruising state and ground takeoff state. The mechanism and quantitative results of slipstream effects on inlet under different conditions, and differences of CRP aerodynamic performance before and after coupled inlet were summarized. Results indicate that the hub vortex produced by the CRP is the primary factor influencing the flow characteristics within the intake duct. Under the cruise condition, wake vortices of the rear row dominated the vortex structure of the inlet and cause non-uniform flowfield. Under the takeoff condition, the coupling of wake vortices of both front and rear rows was enhanced, leading to the presence of vortices in pairs. These paired vortices collectively impacted the flow field within the inlet. The periodic movement of wake vortices within the inlet induced periodic fluctuations in aerodynamic parameters. When propfan wake vortices dominated the vortex structure, the amplitude of parameter oscillations diminished along the duct, yet the oscillation frequency remained constant and the phase angle maintained a uniform lag. Following the detachment of wall-separated vortices, the duct's vortex structure underwent changes, disrupting the amplitude, frequency, and phase angle of parameter oscillations due to the mixing of these separated vortices. The slipstream enhanced the total pressure of inlet; however, intricate trailing vortices significantly exacerbated the total pressure and swirl distortions at the duct's exit, leading to pronounced uneven internal flow distributions. Internal flow losses increased by 2.85 times during cruising and 1.09 times during takeoff. In addition, integration of the intake duct altered the internal flow field structure of the rear blade channel of the counter-rotating propfan. This results in a decrease in velocity within the channel, a forward shift in the position of the shock wave on the blade suction surface, and a modification of the operational characteristics of blade elements at varying heights. Consequently, this improved the thrust performance of the propeller fan. Under the design condition, the thrust coefficient of the front row blades rose by 5.9%, that of the rear row blades increased by 27%, and the overall efficiency of the propfan improved by 5.2%.

To investigate the infrared radiation characteristics of plumes of hypersonic vehicles in near space, a computational methodology for the flow field of the near space plume infrared radiation was developed. Initially, the numerical simulation of the plume flow field is carried out by using the nonlinear coupling constitutive relationship model, and the accuracy of the flow field calculation model is verified. Subsequently, a comprehensive dataset consisting of 46 200 sets of flow field characteristic parameters and absorption coefficients was generated using a line-by-line calculation physical model, informed by a full-factorial experimental design. This dataset was employed to train an accelerated line-by-line calculation model based on a BP neural network. The trained model demonstrated a maximum mean absolute error of 0.003 65 and an R² value of 0.999 4, achieving computational speeds four times faster than the traditional line-by-line calculation physical model. Finally, combined with the Backward Monte Carlo method as the infrared radiation transmission method, the infrared radiation calculation method of the hypersonic vehicles plume flow field is established. The infrared radiation characteristics of the plume under the cruising state of X-51A aircraft were studied by using the established flow field and infrared radiation calculation methods. The findings indicate that hypersonic plumes in near space exhibit a phenomenon where external heat flow envelops a cooler internal flow. Furthermore, the infrared radiation is stronger in the stagnation region at the plume's periphery, resulting in a “scissors” pattern in infrared imaging. Variations in the shear layer significantly influence the infrared imaging characteristics of the plume.

To efficiently calculate the infrared radiation characteristics of serpentine nozzles with varying shielding properties during the early stages of infrared stealth design for exhaust systems, a rapid prediction model for the tail infrared radiation characteristics of serpentine nozzles has been developed. This model combines the image-based method and the one-dimensional flow field model. The former transforms the geometric shielding relationship of serpentine nozzles into pixel operations on multi-layer two-dimensional images, and the latter is established based on the compressible pipe flow and jet theory. The gas absorption and emission characteristics are calculated using the statistical narrow-band model. The impact of shielding properties on gas radiation is also considered by fitting functions for the length and volume of the gas-visible region. The computational efficiency and accuracy of the model are verified by comparison with the discrete transfer method. The results show that the rapid prediction model can shorten the computation time of the tail infrared radiation from several hours to about 1 s, and the outcomes of the model align well with the discrete transfer method in terms of trend and value. The maximum relative error in total infrared radiation intensity for different shielding properties of serpentine nozzles and various nozzle working conditions is only 6.5%. Furthermore, the maximum relative errors for wall radiation and gas radiation are 4.1% and 5.2%, respectively. The model also demonstrates good generalizability for axisymmetric and two-dimensional nozzles.