With the rapid development of infrared detection and guidance technology, future fighter aircraft will face the search and detection of enemy aerial, air and ground all-round infrared detection systems. To determine the average infrared radiation intensity in the rear hemisphere of straight flow exhaust systems, an approximate calculation method was designed basing on limited data. This method was applied to conduct numerical simulations of the average infrared radiation intensity in the rear hemisphere for axisymmetric nozzles, two-dimensional convergent-divergent rectangular nozzles, and two-dimensional ejector nozzles. The innovative aspect of this method is that it can obtain the distribution of infrared radiation intensity in the rear hemisphere and the average infrared radiation intensity for the corresponding exhaust system using infrared radiation intensity data from typical detection planes, thereby significantly reducing calculation time. The results demonstrated that with this newly developed approximate calculation method, the error in the distribution of the average infrared radiation intensity for each model is generally within 10%, and the error in the average infrared radiation intensity is around 5%. These simulation calculation errors are acceptable for engineering applications.

Brush seals are advanced seal widely used in aero-engines. The friction and wear characteristics of the brush seal have great influence on the performance and life of the brush seal. In this paper, the ball-disk friction and wear experiment has been proposed to measure the friction and wear properties of the brush seal material. It is noteworthy that the ball-disk friction and wear experiment could measure friction coefficient and wear factor accurately, only if the working conditions are reasonably set. The experimental results demonstrate that the temperature significantly affects the friction and wear properties of the brush seal, however, the load does not within a specific range. The ball-disk friction and wear experiment is a good choice for its lower cost, simpler system and acceptable accuracy. The experimental results provide important basic data for the rational selection of the candidate materials.

Abstract Low hysteresis brush seals are frequently used in systems operating at a high speed in environments featuring a high temperature and pressure. The high-speed rotor comes into contact with the bottom end of the bristles of the brush seal in such scenarios to generate a significant amount of frictional heat, which directly affects its sealing performance and service life, while a high temperature can increase the magnitude of its friction-induced heat. In this study, we report the cyclical testing of a low hysteresis brush seal while increasing and decreasing the speed of the rotor under varying differences in the pressure and temperature. We focus on the characteristics of leakage, hysteresis effect, and temperature rise at the bottom end of the bristles of the brush seal owing to frictional heat. The results showed that the volume of leakage increased at a high temperature under a strong hysteresis effect and a small rise in the temperature. Moreover, as the speed of the rotor exceeded 6000 rpm, the temperature of the system increased rapidly owing to frictional heat. During the speed decrease stage, the temperature rise decreased sharply and then gradually until it became nearly constant. The hysteresis effect resulted in a lower temperature rise during the decrease in the speed of the rotor compared with that during an increase in its speed. While the low hysteresis structure can effectively reduce hysteresis effect compared with that of the conventional brush seal, it induced greater leakage. It is necessary to choose a pressure relief chamber of a suitable size to minimize leakage. Furthermore, the low hysteresis brush seal exhibited a smaller friction temperature rise than the conventional seal, where this is beneficial for its service life.

In this study, we propose a method to estimate the characteristics of infrared radiation of the exhaust system during the design of a turbofan engine based on the overall design parameters in the context of infrared stealth. We initially establish a model to predict the distribution of the velocity and temperature fields, as well as the concentration of gas inside the cavity of the exhaust system and the plume based on the parameters of temperature and pressure of the inner and outer mixing sections of the culvert. Following this, we develop models to predict the intensity of infrared radiation of the plume perpendicular to the axis of the nozzle (alpha = 90 degrees) along with the exhaust system along the axis of the nozzle (alpha = 0 degrees). A comparison between the predictions of the proposed model and the results of numerical calculations showed that the error was no larger than 7 %. Another comparison between the predictions of the proposed model and experimental data showed that the error was little larger than 11 %. The proposed model can thus be used to predict the characteristics of infrared radiation of the turbofan engine in the design stage.

The serpentine convergent-divergent nozzle represents an optimal configuration for next-generation fighter aircraft characterized by low detectability and high thrust-to-weight ratio. In contrast to the serpentine convergent nozzle, such configuration offers increased design flexibility with additional parameters, leading to heightened interactions among these parameters. As such, it is crucial to reveal the influence of design parameters on the aerodynamic performance of the serpentine convergent-divergent nozzle and the multifactor interaction, as well as its mechanism. Therefore, the influence, interaction and sensitivity of parameters on the aerodynamic performance of the nozzle were numerically investigated using the orthogonal test method. Additionally, the influence mechanism of the convergence angle, throat aspect ratio, and axial length to inlet diameter on the flow characteristics of the nozzle was investigated in detail. The results show that the convergence angle is identified as the main factor affecting the aerodynamic parameters of the nozzle. As the convergence angle increases, the thrust coefficient, total pressure recovery coefficient and discharge coefficient gradually decrease. The interaction between throat aspect ratio and other parameters is obvious. Different design parameters affect the local loss and the friction loss by affecting the curvature and wetted perimeter area, resulting in different aerodynamic characteristics of serpentine convergent-divergent nozzle.

It is widely agreed that the infrared stealth should be considered in the preliminary design phase of turbofan engine so as to decrease the infrared radiation of engine exhaust system. In the present study, a turbofan engine with afterburner at small bypass ratio (0.44, 0.5) in the flight envelope (altitude h < 12.8 km, Mach number M < 1.5) was studied, in which afterburner state was not taken into account. The design method of geometrical model of the exhaust system at the preliminary design phase was developed, and the Latin hypercube sampling method was used to design sample cases, which were numerically simulated to study the infrared radiation characteristics of the exhaust system at different inlet parameters of the exhaust system. A neural network model was developed to predict the infrared radiation of the exhaust system, in which the inlet parameters of the exhaust system in the flight envelope were applied as input parameters. Results show that the maximum error of the developed neural network model for predicting infrared radiation characteristics is less than 10 %, which can be used to evaluate the infrared radiation characteristics of the exhaust system at the preliminary design phase of turbofan engine.

Based on the orthogonal experimental method, a simulation case of the flow field of the ejector nozzle was designed to investigate the influence of the structural parameters of the ejector nozzle on the internal and external flow. This study explored the effects of throat area, outlet area, throat position, and ejector nozzle length on the ejector flow rate ratio, thrust coefficient, and net thrust coefficient. Subsequently, flow path geometry optimization was conducted to maximize the thrust coefficient or net thrust coefficient. The results revealed that the throat area ratio and the outlet area of the ejector nozzle are the primary factors affecting the aerodynamic performance. Compared to the baseline ejector nozzle model, the optimal model for thrust coefficient exhibited a 16.333% improvement, while the optimal model for net thrust coefficient demonstrated a significant enhancement of 46.674%.

Installing a precooler behind the intake is an effective approach for hypersonic air-breathing pre-cooled engine to cool the hot incoming air. Synergetic air-breathing rocket engine is a revolutionary hypersonic air-breathing pre-cooled engine with complex thermodynamic cycle i.e. air cycle, helium cycle. Air/helium precooler is a key component and its configuration and operating condition have great effect on the performance characteristics of air-breathing pre-cooled engine. Thus, the minimum periodic flow and heat transfer model of the precooler are established. The effects of key parameters on the heat transfer performance of precooler are numerically studied. The results indicate that: when the tube row number increases from 7 to 15, the average heat transfer coefficient of air side decreases by 57%, the heat exchange rate increases by 19%, and effectiveness increases by 18.4%. The tube transverse pitch can enhance the heat transfer coefficient of air and helium side, while the heat exchange rate decreases by 33 % when the tube transverse pitch increases from 1.5 to 3.5. The helium inlet velocity can improve the heat transfer performance of precooler and reduce the flow resistance of air side.

In order to study the infrared radiation characteristics of an air-breathing hypersonic vehicle powered by a scramjet, it is necessary to solve the internal and external flow field of the air-breathing hypersonic vehicle. Owing to the complexity and difficulty of solving the three-dimensional flow and heat-transfer process in a scramjet combustor, a quasi-one-dimensional calculation method was established. Utilizing zooming technology, a combination of quasi-one-dimensional simulation within the combustion chamber and three-dimensional numerical simulation elsewhere on the vehicle was employed to obtain the flow field. The accuracy of the zooming method in determining flow, heat transfer, and infrared radiation was verified through comparison with experimental data. The results show that under the flight condition of Ma = 6, the gas temperature and wall heat flux in the scramjet combustor first increased and then decreased along the flow direction. The Mach number of the plume was smaller than that of the free flow, while the velocity of the plume was slightly larger. In the wavelength range of 3–5 μm, as the azimuth angle increased, the integrated radiation intensity of the air-breathing hypersonic vehicle demonstrated a characteristic pear-shaped distribution.

Approximate computational methods were developed to determine the infrared radiation (IR) distribution and average IR intensity in the rear hemisphere of the serpentine two-dimensional (2D) exhaust system. These methods use data from horizontal and vertical detection planes. The methods were verified using the IR intensity data on the typical detection planes of the serpentine 2D convergent and convergent–divergent exhaust systems, respectively. The results demonstrate that the methods developed in this paper for calculating the distribution of IR intensity and the average IR intensity of the serpentine 2D exhaust system in the rear hemisphere are feasible. An approximate evaluation of the infrared stealth performance of the exhaust system can be conducted. According to the validation model presented in this study, the calculation error of the average IR intensity of serpentine 2D exhaust systems in the rear hemisphere is less than 5

The precooler is a key component of hypersonic precooled engine, and its actual structural have a high anisotropy in the axial, circumferential, and radial direction. Considering the non-uniform flow inside the precooler is crucial to study the flow and heat transfer characteristics of the precooler. The previous simplified method for the precooler has not fully considered the flow non-uniformity of the axial, radial and circumferential direction. In this paper, a simplified numerical method based on anisotropic porous media model for the precooler was proposed. The unique aspect of this method is the integration of anisotropic surface porosity of porous media model with the zero-dimensional heat transfer model. In addition, the influence of key structural parameters on the flow and heat transfer characteristics of precooler was studied. It is found that as the tube transverse pitch increases from 2 to 3.5, the total flow rate increases by 11%, and the heat exchange rate decreases by 25%. As the number of tube rows increases from 7 to 13, the heat exchange rate increases by 22%, but the total pressure loss coefficient increases from 0.046 to 0.082. In addition, the decrease of inner diameter of the precooler improves the heat exchange rate, but reduces the flow capacity of the precooler. These results indicate that the simplified numerical method is capable of simulating the behavior of precooler and provides some guidance on the design of precooler.

Because of their good sealing performance, brush seals are used widely in turbomachinery as alternatives to traditional labyrinth seals. However, obstructed by friction, the flexible bristles cannot recover fully after deformation, resulting in a hysteresis effect and decreased sealing performance. In the present study, a three-dimensional multi-row bristle model is established that considers the friction between the neighbouring bristles and back plate. Then the deformation and recovery of the bristles during the radial reciprocating offset of the rotor in the presence of a pressure difference are studied. To evaluate the hysteresis effect, analysis is directed at the unrecovered bristle deformation in the radial direction affected by bristle diameter, bristle lay angle, fence height and width of pressure relief chamber. The results show that the larger the bristle diameter and fence height, the weaker the hysteresis effect and the larger the bristle lay angle, the stronger the hysteresis effect. With increasing width of the pressure relief chamber, the hysteresis effect weakens and then strengthens, meaning that the pressure relief chamber has an optimal axial width that departure from enhances the hysteresis effect. The present results offer a reference for selecting the structural parameters of brush seals in engineering design.

The rescaling–recycling method (RRM) is usually used to generate turbulent inflow for the LES of compressible wall-bounded flows, which can lead to relatively high computational cost for high Re flows since the mesh resolution increases exponentially with Re number. A turbulent inflow generation method based on the scaling of low Re flow, referred as TIG-LowRe, is proposed, aiming at reducing the computational cost when applying the RRM. To validate the proposed method, the TIG-LowRe method was applied to generate turbulent inflow for the LES of a non-isothermal round jet flow at Re = 86,000. Two cases were carried out with the inflow generated based on two round pipe flows at Re = 10,000 and 24,000. The results show that the mean and fluctuating temperatures of the two cases agree well with the experimental data. In the case of low Re flow at Re = 10,000, the jet flow decays too fast along the axial direction, the mean and fluctuating axial velocities are over-predicted and the radial fluctuating velocity is under-predicted. By increasing the Re of the low Re flow to 24,000, the decay rate of the jet flow decreases and the accuracies of the mean and fluctuating axial velocities are obviously improved, while the radial fluctuating velocity shifts further away from the experimental data. The main reason for the difference between the two cases is that more fine turbulent structure of the inflow in case-Re10000 is lost than in case-Re24000 during the turbulence generation process.

In order to investigate the leakage and heat transfer of eccentric labyrinth seals, the flow and heat transfer process of eccentric and rotating straight-through labyrinth seals were numerically simulated. The research results show that the influence of eccentricity and rotation on the leakage coefficient is very small, and the change range is less than 2 %. Under the condition of non-eccentricity, the influence of rotating speed on the leakage coefficient depends on the rotating Taylor number. When the Taylor number is greater than 34.2, the leakage coefficient decreases slightly with the increase of rotating speed. At the same pressure ratio and rotating speed, the average Nusselt number of stator is greater than that of rotor. With the increase of eccentricty, the average Nusselt number of the section with larger clearance increases gradually, while the average Nusselt number of the section with smaller clearance decreases gradually.

针对指尖密封提出泄漏流动和固体变形的三维双向耦合计算方法.该方法采用多孔介质模型计算密封片组泄漏流场,将流场计算得到的压力分布加载到有限元模型上进行密封的固体变形计算,再根据变形计算结果更新多孔介质模型参数,重新计算泄漏流动;如此重复,直至泄漏流动和固体变形计算均达到收敛.采用该计算方法,对不同压差下指尖密封的泄漏流动与固体变形进行计算.计算结果表明:流动与变形的耦合计算经过3次迭代达到收敛;指尖密封结构的压降、泄漏及变形主要发生在指尖靴区域,密封片组的最大变形主要取决于径向变形;当上下游压差从0.1 MPa增大至0.5 MPa,密封片的最大径向变形增大约2％,最大轴向变形增大约400％,泄漏量增大约130％.

准确高效的电磁散射仿真方法对设计隐身航空发动机排气系统非常重要.将特征基函数法(CBFM)、多层快速多极子算法(MLFMA)、插值分解(ID)算法和并行技术相结合,对发动机排气系统的单站雷达散射截面(RCS)进行仿真计算.插值分解(ID)可以对单站激励矩阵进行低秩压缩,因此可以减少矩阵方程的求解次数,显著提高了传统CBFM-MLFMA的计算效率.为了验证算法的正确性,对轴对称排气系统模型进行了加工和RCS试验测试.与测试结果相比,仿真结果与其吻合良好,验证了算法的精度.

25 models are designed according to the orthogonal table, to study the comprehensive relationship between heat transfer and sealing characteristics in straight-through labyrinth seals with different geometrical parameters. The experimental investigations are carried out at different pressure ratios of 1.2, 1.3, 1.4, 1.6, 1.8. The results show that different geometry parameters brought obvious different influence on both leakage efficiency and heat transfer characteristics, and in which the parameters (H, c,β) have the most significant influence on leakage efficiency, and heat transfer characteristics under different pressure ratios, whose small changes in size can lead to performance optimization in leakage and heat transfer. The geometrical parameters always show consistent influence on leakage efficiency and heat transfer characteristics, while heat transfer characteristics show larger vibration with the increase of pressure ratio.

The flow fields and infrared signatures of a turbofan engine axisymmetric exhaust system with or without afterburning have been numerically investigated. The flow fields of the exhaust system have been calculated using a commercial numerical simulation software, whereas the infrared signatures have been calculated by the software nuaa-ir (numerical analysis of aircraft's infrared signature) developed by the current authors' research team. The spectral and integral radiation intensity in the waveband of 3-5 & mu;m of the axisymmetric exhaust system have been calculated and analyzed. The results show that the thrust of the axisymmetric exhaust system with afterburning is increased by 55% as compared with nonafterburning. The main infrared radiation source of the exhaust system with afterburning is the plume. The infrared radiation of the plume accounts for 65% of the total infrared radiation of the exhaust system in the downstream axial direction. The maximum integral radiation intensity of the exhaust system with afterburning is nearly seventeen times that of without afterburning. The integral radiation intensity of the exhaust system in the downstream axial direction with afterburning is about four times that of without afterburning. The integral radiation intensity of the exhaust system in the lateral direction is about 54 times higher than that without afterburning.

横向射流是降低航空发动机尾喷流在3～5μm波段上红外辐射特征的一种高效低阻的途径.本文以收敛喷管为对象,在喷口下游布置了一对上下对称的扁平横向射流孔,数值模拟研究了两股同频横向射流之间相位差(0°,90°和180°)对下游掺混与红外抑制的影响规律.结果表明,脉动横向射流诱导产生的流向涡强化了掺混,不同横向射流之间的相位差会使涡对出现的时间与空间复杂交错,可以充分卷吸气流,加强尾喷流与外界大气的掺混.而相位差的优化设计可以进一步提高涡对之间的掺混增益,增强红外抑制效果的作用.与Δφ=0无相位差状态相比,随着相位差增大,红外抑制效果逐渐增强.相位差为π时,掺混效果最好,高温区长度降低约14％,红外积分辐射特征最大降幅约13％,光谱辐射降幅最大可达21％,有效增强了对尾喷流的红外抑制.