During the ascent phase of the Mars Ascent Vehicle (MAV), the thermal environment of the launch system becomes highly complex and severe due to plume-wall interactions and the unique characteristics of the Martian atmosphere. This study employs a Reynolds-Averaged computational fluid dynamics methodology combined with an overset grid technique to simulate the transient nozzle flow field evolution throughout the MAV ascent phase. By employing finite-rate chemical kinetics, a model was established to simulate the complex dissociation-recombination reactions between the high-temperature propellants of MAV and the Martian atmosphere. To demonstrate the necessity of accounting for finite-rate chemistry, computational results of reacting and non-reacting flows were compared. Furthermore, variations in the thermal environment of the launch system during the ascent phase were investigated using the discrete ordinates method. Results reveal that the inclusion of chemical reactions leads to an 11.2% increase in peak temperature during ascent due to plume-wall impingement. As the MAV ascends, significant changes occur in the heating rate and pressure distribution across various components of the launch system. The surface pressure at the MAV base mount reaches its maximum around 0.4 s, exceeding values at other time instances by 15.2-30.7%, while the heat flux at the base peaks at 81.2 kW/m2 at 0.42 s. The launch pad, elevation mechanism of launch, and launch pad support attain their peak heat flux values of 544 kW/m2, 29.8 kW/m2, and 3240 kW/m2 at 0.22 s, 0.32 s, and 0.23 s, respectively. The peak heat flux for each component is 20-200% higher than that at other moments. With the exception of the elevation mechanism of launch, convective heat flux constitutes the dominant mode of surface heat transfer for all components. (c) 2026 Published by Elsevier B.V. on behalf of COSPAR.

This study investigates the complex flow evolution and thermal load distribution induced by multi-nozzle gas jet impingement on platforms during heavy-lift launch vehicle recovery. A high-fidelity numerical model is developed based on the three-dimensional compressible Navier–Stokes equations and the realizable k-ε turbulence model. A structured hexahedral grid is used to discretize the computational domain, and the numerical method is validated against Planar Laser-Induced Fluorescence (PLIF) experimental data for under-expanded impinging jets. The effects of descent altitude, horizontal drift distance, and protective wall height on flow characteristics and thermal environment are systematically examined. The descent process can be divided into three stages: high-altitude free expansion, transitional impingement, and near-ground wall-bounded flow. As the altitude decreases, the flow structure evolves from regular Mach disks to fragmented shock waves and large-area wall-attached jets, accompanied by a significant expansion of the high-temperature region on the platform surface. Horizontal drift leads to two distinct physical mechanisms that undermine the protective effect of the wall. Under moderate relative drift, this mechanism manifests as a vortex-overtopping mode, in which the shear layer at the wall top destabilizes and forms a cascade of vortices. This establishes a steady cross-wall energy transport pathway, substantially increasing the peak heat flux in the equipment zone. Under large relative drift, a shear-dispersion mode occurs, reducing the peak heat flux by over 80% and shifting the threat from concentrated transport to dispersed dissipation. These two mechanisms can be unified by the drift ratio η=S/H (H is the landing altitude and S is the horizontal drift distance). For a wall height of 1 m, the critical value is ηcrit=0.27. When η is below this value, the wall effectively blocks the high-temperature gas; when η exceeds this value, the protective function is compromised. The parameter H‾w is defined as the critical dimensionless protective wall height. In the present configuration, H‾w=2.56, which corresponds to a physical height of 3.5 m. At this critical state, the peak heat flux for the vortex-overtopping mode is reduced by 99.1% compared with the failure condition. Consequently, the cross-wall energy transport pathway is effectively suppressed, and the protection mechanism shifts fundamentally from magnitude suppression to structural restriction of the flow field. Overall, this study reveals the physical mechanisms of protection failure under horizontal drift, establishes a dimensionless failure criterion based on the relative drift ratio, and proposes a specific dimensionless critical wall height, offering engineering guidance for the design of thermal protection systems for reusable launch vehicle recovery platforms.

To investigate the plume flowfield and thermal environment of heavy-lift launch vehicles, a numerical model for a nine-nozzle launch vehicle was established based on the three-dimensional compressible Navier-Stokes equations, the Realizable k-epsilon turbulence model, the DOM radiation model, and a finite-rate chemical kinetics model. The study indicates that as the rocket ascends, the compressive effect of the external environment on the exhaust plumes gradually diminishes. Multiple plumes collide to form collision zones, and following these collisions, a gas flow directed towards the rocket base gradually forms. At altitudes below 10 km, radiation heat flux dominates the bottom heat transfer. As altitude increases, convection heat flux becomes predominant, with the total bottom heat flux peaking at 1273.70 kW/m2 at 30 km. At low altitudes, the temperature difference between the frozen flow and reaction flow is more pronounced. At 20 km altitude, the peak temperature in the reaction flow plume field is 8.28% higher than that in the frozen flow. At high altitudes, the temperature difference between the frozen flow and the reaction flow is not significant. At an altitude of 40 km, the peak temperature of the reaction flow exhaust plume is 7.15% higher than that of the frozen flow.

During high-altitude Mars Ascent Vehicle (MAV) flight, the base thermal environment becomes highly complex and severe due to high-speed freestream flow and nozzle deflection for attitude adjustment. Finite-rate chemistry kinetics was employed to compute combustion reactions, and a computational model for predicting the MAV base thermal environment at high altitude was developed based on the Discrete Ordinates Method (DOM). Computational results were compared with wind tunnel experimental data. The base thermal environment variations under three nozzle deflection angles across six distinct flight altitudes were simulated. The results indicate that the effect of chemical reactions increases the peak temperature in the MAV exhaust plume by 3-6 %, and the Mach number decreases. Accounting for chemical reactions leads to an approximate 2.5 % increase in base heat flux at different heights. When the engine works normally, the MAV base heat flux initially increases then decreases with rising flight altitude, reaching a maximum value of 22.1 kW/m2 at 16.8 km. The larger nozzle deflection angles result in higher base heat flux. However, the increment in base heat flux induced by engine deflection diminishes as altitude increases. When the flight altitude increases from 4.2 km to 34.6 km, the increment in base heat flux for a 10 degrees nozzle deflection decreases from 13.9 % to 7.8 % compared with a deflection of 0 degrees. The highest base heat flux of 23.8 kW/m2 occurs at 16.8 km with a 10 degrees nozzle deflection.

Violent shaking induced by motion excitation in a dynamic environment can generate significant additional forces and moments in liquid propellants, affecting spacecraft stability and attitude control. This study employs the CEL method to establish a fluid-structure interaction model for a launch vehicle fuel tank. The penalty function approach addresses FSI and hinged contact behavior, while the control volume method quantifies liquid pressure on anti-sloshing plates and propellant volume changes. Initial filling angles were used to simulate motion excitation equivalently. For partial filling conditions, the effects of equivalent filling height, initial filling angle, and anti-shake plate thickness on impact resistance were analyzed. Results indicate that at a fixed initial liquid filling angle, increasing the equivalent liquid filling height elevates both the maximum pressure and displacement of the anti-sloshing plate, though the rate of increase diminishes with greater plate thickness. At a fixed equivalent liquid filling height, increasing the initial liquid filling angle raises the maximum pressure, while the enlarged liquid-solid contact volume suppresses propellant sloshing, thereby reducing the displacement of the anti-sloshing plate. This study provides theoretical support for the structural design of anti-sloshing plates in rocket tanks.

The flow field during launch vehicle takeoff and the thermal environment of the launch platform are investigated by numerical simulation. A computational model is constructed based on the Realizable k-ε two-equation model and the convective heat transfer model proposed by Launder. The results show that with the increase of the rocket takeoff height and the increase of the rocket drift, the impact of the gas jet on the launch platform increases and then decreases, and the impact is the smallest when the takeoff height is 2m and the largest when it is 20m.

This paper investigates the heat distribution on the movable vertical arm of the CZ-12 launch vehicle within the rocket plume impact field in the three-horizontal test launch mode. A model for the different flight altitudes of rocket plume impact on the different angles of the vertical arm was established based on the three-dimensional Navier–Stokes equations and a realizable k−ε turbulence model. The numerical results were compared with experimental data and schlieren images from literature to verify the effectiveness and accuracy of the established numerical model. The results show that when the flight altitude of the rocket is between 30 m and 40 m, the worst heat environment occurs on the front and bottom of the vertical arm. Before reaching a flight altitude of 30 m, a smaller rotation angle of the vertical arm leads to higher maximum temperatures at these two regions. After reaching a flight altitude of 40 m, a larger rotation angle of the vertical arm results in higher maximum temperatures. The top of the lower frame structure is not directly affected by the rocket plume before reaching a flight altitude of 30 m. After reaching a flight altitude of 40 m, a smaller rotation angle of the vertical arm results in higher heat loads on the frame. The results of this study can provide a basis for designing targeted thermal protection for vertical arms. They also contribute a new idea for reducing the thermal load on the vertical arm, which is to rotate the vertical arm to the appropriate angle according to the rocket takeoff altitude. Meanwhile, these research findings will supply a relative reference for researchers who are concerned about other facilities in the surrounding area.

Numerical simulation of an underwater single-tube and multi-tube low-speed bubble flow exhaust process and acoustic signals based on the fluid volume model (VOF) and Williams Hawkins (FW-H) model is conducted to study the acoustic signals and noise reduction mechanism of multi-tube low-speed bubble flow exhaust process. Compare the simulation results with a theoretical basis to verify the applicability and accuracy of the model. The research results indicate that multi-tube exhaust can significantly reduce pressure fluctuations caused by bubble detachment from the pipe mouth, and can reduce the amplitude of noise in the mid to low frequency range.

The launch vehicle's second stage spends most of its time at high altitudes, resulting in extremely complex flow conditions at the base of the rocket due to the simultaneous operation of the main nozzle, the swing nozzles, and the exhaust pipes. To deeply study the plume flow field and thermal environment of the launch vehicle's second stage, a supersonic exhaust plume model is established by using the three-dimensional Reynolds-averaged Navier-Stokes (RANS) method and the realizable k-e turbulence model, and the radiation model is based on the discrete coordinate method (DOM). It is shown that with the increasing flight altitude of the rocket, the compression of the nozzle jet by the forward incoming flow weakens, the barrel-shaped compression wave in the exhaust plume increases, and the nozzle jet can only be fully expanded in the region farther downstream of the rocket. The higher temperature at the rocket head is due to the decrease in flow velocity and increase in pressure and temperature after the supersonic airflow forms a detached shock wave, and the airflow at the rocket head reaches a stagnant state. At high altitudes, the total heat flux in the rocket base is dominated by the radiation heat flux, which represents more than 72.57 % of the total heat flux, reaching 83.47 % at the flight altitude of 65 km. The total heat flux in the sidewall of the rocket is dominated by the radiation heat flux, which represents more than 67.26 % of the total heat flux. (c) 2025 COSPAR. Published by Elsevier B.V. All rights are reserved, including those for text and data mining, AI training, and similar technologies.

This paper investigates the surface thermal environment of a rail-based launch system subjected to missile plume impingement flow during hot launch. This study established a computational model for missile plume impingement on a rail-based launch system based on the three-dimensional Navier-Stokes equations and the realizable k-epsilon turbulence model. CFD simulations were performed for the flow field impacting the launch system with varying deflector heights under different missile flight altitudes. The results demonstrate that when the missile flight altitude H is within 20 m, the launch system experiences a severe thermal environment, the maximum gas temperature on the deflector surface can reach as high as 3200 K, the maximum gas temperature on the surface of the carriage bottom will also exceed 2600 K. Higher deflector heights improve the thermal conditions for facilities beneath the launch vehicle, such as the rail track components and sleepers, it can reduce the maximum surface temperature of the carriage bottom by up to 22.3%, but simultaneously deteriorate the thermal environment on the upper surface of the launch vehicle and the deflector itself. Furthermore, the position where the barrel shock of the engine plume impinges on the deflector alters the gas temperature distribution pattern on the deflector surface. This demonstrates that even a slight variation in the engine's position relative to the deflector can induce dramatic changes in the gas temperature distribution morphology across the deflector surface. Research demonstrates that during rail-based launch system operations, employing deflectors with optimized heights can significantly improve the thermal environment across critical components. For deflectors of a given height, the current engineering practice of using discrete computational conditions (e.g., H = 0 m, 2 m, and 10 m) requires finer parametric refinement. This is essential to resolve the phenomenon where minor variations in engine-deflector standoff distance induce significant morphological changes in surface gas temperature distribution, thereby enabling further optimization of the launch system's thermal protection design. The "thermal environment" in this paper only provides the surface gas temperature as a reference.

In response to the influence of different gas models on the flow field structure and thermal environment of a single nozzle rocket first-stage in retro-propulsion, a two-dimensional axisymmetric N-S equation containing chemical reactions was numerically solved. The reverse jet flow field of the rocket configuration was numerically simulated using a frozen perfect gas model and a reacting gas model, and the changes in surface heat flux along the trajectory point of the rocket were analyzed. Research has found that chemical reaction gas models can more accurately describe energy exchange and component changes in flow fields. For the thermal environment of the rocket body, especially at the bottom and top of the rocket body, there is a significant difference between the peak heat flux calculated by the chemical reaction flow and the perfect gas model. Therefore, to accurately depict the supersonic retro-propulsion flow field and aerodynamic thermal properties of the first-stage rocket’s deceleration, advanced measurement and analysis methods are essential, ensuring a precise understanding of flow dynamics and thermal behaviors.

In this paper, the computational fluid dynamics (CFD) simulation of the impact flow field of the gas flow in the railway launch system is carried out. Based on 3D Navier Stokes Equation and realizable turbulence model, the model of plume impact on Railway launch system is established, and the plume impact flow field of missile at different flight altitudes is analyzed. The results show that the deflector can effectively protect the rail and sleeper. The impact on the deflector is the most severe when the flight altitude is 0 meter, and the thermal load of the deflector will be greatly reduced when the flight altitude is more than 10 meters. The sharp increase of pressure caused by the deflector located at the Mach disk node at a specific height will not cause the sharp increase of temperature to the same extent. The train compartment is not directly under the rocket engine, so it will be impacted when the missile flies at a higher altitude, but the peak temperature and pressure are far less than those at the deflector. The research shows that when the railway launch system works, the deflector is needed to protect the railway infrastructure, and the corresponding thermal protection work can be carried out according to the thermal load characteristics of the train carriage.

Abstract This study aims to investigate how afterburning affects the thermal environment of the launch pad during a rocket launch. The thermal environment and launch pad flow field are studied using numerical simulations during the vehicle lift-off phase to achieve this goal. The launch vehicle exhaust plume model is constructed during the lift-off phase using the reaction model, the Realizable k-ε two-equation turbulence model, and the three-dimensional compressible Navier-Stokes equations. The results of the investigation demonstrate that the temperature field of the rocket exhaust plume is higher in the gas chemical reaction flow scenario compared to the frozen flow scenario. In addition, considering both the frozen and gas chemical reaction flows causes the launch pad surface temperature to rise. The research approach used in this study sheds important light on the launch pad’s thermal protection design.

This research investigates the water spray on the deflector for the cooling effect of the launch platform and compares the temperature distribution on the deflector’s bottom surface at various water spraying rates. A one-nozzle launch vehicle gas jet model with drift was simulated using the three-dimensional compressible Navier–Stokes equations, the realizable [Formula: see text] turbulence model, and the Eulerian discrete phase (EDP) model. The installation of the water spray cooling device was demonstrated to reduce the area of the deflector’s high-temperature region because water accumulates at its bottom surface, acting as a protective layer. This causes the high-temperature gas jet to contact the water instead of directly impacting the deflector, while the water evaporates and absorbs heat. Compared with conditions without the water spray cooling device, temperatures across most areas of the deflector’s bottom surface were reduced by approximately 76%. As water spray velocity increases, the peak temperature and the high-temperature region on the deflector’s bottom surface gradually decrease, and the temperature in the area directly impacted by rocket exhaust also decreases. At a water spray velocity of [Formula: see text], the lowest peak temperature occurs on the deflector’s bottom surface, representing a reduction of up to 37.6% compared to that at [Formula: see text].

This paper studies the one-nozzle liquid rocket exhaust plume impinging on the flame deflector. The effect of drift and afterburning on exhaust plume impinging on the deflector during the takeoff phase is studied by numerical simulation. An impact model is established based on three-dimensional compressible Navier-Stokes equations, a two-equation realizable k-epsilon turbulence model, an afterburning model, and second-order total variation diminishing upwind scheme. The study shows that the smoothness of the deflector discharge is affected by the impingement position of the exhaust plume impinging on the deflector due to the drift of such a position during the take-off phase of the rocket. The temperature of the rocket exhaust plume flow field after considering the chemical reactions is higher than that of the frozen flow, and the temperature of the deflector surface is also higher than that of the frozen case. With the increase of the impingement angleat inlet areaof the rocket exhaust plume impinging on the deflector, the deflector discharge performance gradually decreases. At the same impingement angleat inlet area, as the impingement angleat exit areadecreases, the deflector slot discharges more smoothly.

During the supersonic re-entry of multi-nozzle heavy rockets into the atmosphere, the basic flow state becomes increasingly complex due to the coupling effect between the retro-propulsion plumes and the freestream. A numerical method using the hybrid Reynolds-Averaged Navier-Stokes and Large Eddy Simulation (RES) method and discrete coordinate method is developed to accurately estimate the thermal environment. In addition, finite rate chemical kinetics is used to calculate the afterburning reactions. The numerical results agree well with wind tunnel data, which confirms the validity and accuracy of the numerical method. Computations are conducted for the heavy carrier rocket re-entry from 53.1 km to 39.5 km altitude with 180° angle of attack by using three different Supersonic Retro-Propulsion (SRP) modes. The numerical results reveal that these three SRP flow fields are all Short Penetration Models (SPM). As the re-entry altitudes decrease, both the plume-plume interaction and the plume-freestream interaction become weaker. The highest temperatures in the plume shear layers of the three SRP modes increase by 8.36%, 7.33% and 6.92% respectively after considering afterburning reactions, and all occur at a re-entry altitude of 39.5 km. As the rocket re-enters the atmosphere, the maximum heat flux on the rocket base plate of three SRP modes stabilizes at 290, 170 and 200 kW/m2 respectively, but the maximum heat flux on the side wall increases significantly. When the altitude declines to 39.5 km, the extreme heat flux of the three modes increase by 84.16%, 49.45% and 62.97% respectively compared to that at 53.1 km.

The pressure-time relation in the shock flow field of a near-earth air blast is complex. The triple point (TP) path is the physical boundary between the free and non-free shock flow fields. Accurately predicting the TP path is the basis for studying the evolution law of the reflection flow field and the guarantee of effectively assessing the damage power of the warhead. Based on the assumption that the Mach stem center is on the projection point of the blast center on the ground, the TP path calculation method was constructed by the geometric relationship. The unknown coefficients were solved using the existing TP data and the least square method. The TP prediction model proposed was compared with the existing ones on the calibration, new numerical simulation TP, and the measured real blast datasets. The error of the new numerical simulation TP data is within +/- 15% of the real value. The results show that the TP path prediction model proposed performs better. Most of its prediction results are within +/- 20% of three datasets compared to other models for the working conditions with the scaled height of burst from 0.397 to 2.777 m/kg(1/3) and the horizontal scaled distance within 10 m/kg(1/3) in the conventional cylindrical TNT explosion with the length-diameter of 1 in the air. The reliability of the prediction model is verified.

Abstract Cooling effect of the water jet on the face of the deflector during offshore launching process of the vehicle is studied in this paper. Based on the Reynolds averaging method and three-dimensional Navier-Stokes equations, a realizable k-ε turbulence model simulates the high-speed gas flow in a rocket engine, and the Euler discrete method simulates the gas-liquid coupling between water and engine gas. The accuracy of the above numerical method is verified by comparing the flow field simulation results with the water spray experiment results. Research has shown that during the launch process of the launch vehicle, the surface thermal environment of the 0-meter-high deflector is the most severe. After cooling by spraying water, the cooling of the jet core area on the surface of the deflector is not obvious, but the temperature of other areas decreases significantly. The pressure on the surface of the deflector will not change significantly. Results provide effective solutions for improving thermal environment of the deflector surface during rocket offshore launching and reducing the deflector surface erosion.

为解决某型号发射装置工作过程中箱内冲击波强度不足的问题,采用数值模拟方法,结合动态网格自适应技术,对燃气冲击波开盖技术中涉及到的流动现象和箱内初始冲击波的产生机制进行研究.通过与相关试验结果对比分析验证所提数值方法的精度,以及当前自适应方法在对关心的流场特征,如冲击波阵面、接触间断等,进行动态加密或稀疏时有可靠的性能.以平板型后盖为例(原始设计)详细阐述了箱内冲击波的演化机制,为更好地认识该技术的工作机制提供帮助;进一步对比了穹顶型后盖对箱内冲击波形成的影响.研究结果表明,与平板型后盖相比,采用穹顶型后盖结构后,可影响气流与箱体尾部侧壁撞击时的角度,减缓该位置湍流区域的形成,从而显著提升箱内初始冲击波强度,使作用在前盖上的超压峰值增加约48％.所得研究结果对发射箱总体方案的设计具有重要的理论意义和工程应用价值.