Test equipment was developed to test the dynamic forces on the cable of a parachute tear-band for high-speed impact loads. The mechanics of the test equipment were simplified as a spring damping model with the system dynamics equations then derived based on the second Lagrange equation. The Runge-Kutta method was used to solve the equations to calculate the tension in the cable during the band tearing test. The results showed that the rope damping plays a major role in the early stage of the high-speed impact loading. The influences of the elastic modulus, damping and mass ratio of the two ends on the cable forces were then further studied. Comparison of the theoretical model with experimental data shows that the predicted cable forces are consistent with the experimental data, which verifies the model accuracy. The results of this study can guide the design of tear-band test equipment.

[Objective] The quantity, total mass, and distribution region of space debris are constantly increasing, and simultaneously, over half of the low earth orbit satellite operators have no sustainable way to remove the failed satellites from space. Consequently, the growth rate of space debris after the end of a satellite's life considerably increases over time. Thus, the problem of removing space debris is very important.[Methods] Drag deorbiting is an efficient way to avoid space debris after micro-satellites fail. In this study, aiming at the deorbit problem using an inflatable film drag balloon, a design of omnidirectional drag after the inflated film balloon is presented. First, a hybrid folding method involving a closed 3D spherical shrinkage into a star-shaped compaction is proposed, which realizes a dense cube shape with zero-line width and variable thickness folding. Then, the air resistance effect in the low earth orbit is an important factor affecting the orbital height of satellite debris according to the atmospheric perturbation theory in drag balloon design, and NRLMSIE-00 model is used to predict the orbital atmospheric density. Based on the assumption of a small deformation spherical shell, the ultimate load regarding the spherical instability of the film drag balloon is analyzed, when the balloon is subjected the maximum air resistance at the height of 200 km. The ultimate loads of the film balloon at room (20℃) and high temperatures (80℃) are compared, and a test is validated in a vacuum chamber by a film balloon. Finally, the inflated balloon dragged by the space debris is considered as the perturbed deorbit motion caused by the air resistance effect. The relationship between the surface-to-mass ratio of various space debris and the balloon diameter with the deorbit time is analyzed, as well as the relationship of the deorbit time of the drag balloon with the orbit height.[Results] Results showed that the polyimide film balloon can be used as a design for an omnidirectional drag deorbit for space debris. This hybrid folding method to a closed 3D inflatable sphere was used for the film balloon with a diameter of 1.8 m, which was reduced to 1/6 000 of its original volume after being folded into a dense cube shape. When the inflated film balloon with a deorbited micro-satellite, it was subjected to a small air resistance effect in the low-Earth orbit, the ultimate load of the polyimide film balloon with diameter of 1.8 m and thickness of 12.5 μm was within the safe range.[Conclusions] In brief, the ultimate load decreases exponentially with the increasement of balloon diameter; the larger the film thickness, the greater the ultimate load that the balloon can withstand. The ultimate load at room temperature is 0.400 0 Pa, while that at 80℃ is 0.330 0 Pa, thus, it is reduced by 17.5% with the increase in temperature. Under the same surface-to-mass ratio, the deorbit time increases with the increase of the deorbit height of the debris, however, under the same orbital height, the larger the surface-to-mass ratio of the space debris, the shorter the deorbit time.

[Objective] Ejection and deployment are the first key actions in the working process of a parachute, thereby resulting in its inflation. Ejection and deployment typically work in the wake of space crafts; therefore, the wake greatly influences the process. The conditions of the free stream may affect the wake flow, such as the attitude, Mach number, and angle of attack. Consequently, the angle of attack of an aircraft is an important consideration in parachute design.[Methods] Ejection and deployment are related to multiple factors, including fluid dynamics, multibody separation, and flexibility deformation. Owing to the advantages of fluid field information and low cost, computation fluid dynamics has become a powerful tool for solving engineering problems regarding fluids. This article focuses on the effect of attack angle on the ejection and deployment of a parachute. Based on the overset grid technology, the three-dimensional unsteady Reynolds averaged Navier-Stokes (URANS) coupled with a six-degrees-of-freedom (6DoF) equation of motion is applied to the research. The simulation consists of two steps:the static flow field simulation and the dynamic separation process based on the static flow field.[Results] The simulation results showed the following:(1) The negative aero force at 0° angle of attack in the recirculation zone hindered the separator department. (2) The wake during the 0° angle of attack was parallel to the axis of capsule; however, 10° and 20° angles of attack demonstrated an obvious deviation from the axis. (3) The separator showed almost no attitude variation for 0° angle of attack, whereas an obvious attitude variation for 10° and 20° angles of attack resulted from the direction of the wake. (4) An obvious interaction occurred between the wake behind the capsule and shock before the separator.[Conclusions] The following conclusions can be drawn from the research:The effects of the angle of attack significantly change the characteristics of the wake. Compared with the 0° condition, the capsule wakes present an asymmetric character in the other two conditions. What's more, the wake shows an obvious deviation from the direction of the initial separated velocity; the asymmetrical wake will affect the trajectory and attitude of the separator; and the effect of the angle of attack will change the relative position between the capsule and the separator and influence the separated time:the time of ejection and deployment will be reduced with an increase in the angle of attack. The method and the conclusions can provide a valuable reference for the validation and design of the recovery system.

[Objective] Previous studies on ringsail parachutes based on computational fluid dynamics (CFD) mainly focused on the slotted structures of the canopy or the overall fabric permeability of ringsail parachutes. Ringsail parachutes with different sail structures, including the upward-exhaust ringsail (UER) and downward-exhaust ringsail (DER), are proposed to allow for different exhaust directions. Sail structures usually account for larger areas and are more complex; therefore, they have a more significant influence on the aerodynamics of parachutes. However, systematic research on the sail structures of ringsail parachutes is lacking. Therefore, the effect of sail structural designs on the deceleration performance of ringsail parachutes is unclear.[Method] In this study, the numerical models of the flow field around a parachute with different exhaust sail directions are established according to the aerodynamic shape of a ringsail parachute under steady descent. The influences of the sail exhaust direction and sail structure fullness on aerodynamic performance are explored via CFD. Wake, jet flow, and canopy surface pressure are explored, the deceleration and stability performances of parachutes with different sail exhaust directions are compared, and the action mechanism is analyzed. Sail fullness is varied as the main parameter, and the relationship between sail configurations and ringsail aerodynamics is analyzed.[Results] The results reveal the following:(1) Jet flow existed at crescent slots. In a UER parachute, jet flow converged in the downstream air column, which reduced the drag and generates an additional restoration moment on the parachute. Therefore, the UER featured smaller drag coefficients but better stabilities than the DER. (2) The UER was significantly affected by geometric permeability; therefore, with increasing sail fullness, the drag coefficient of the UER increased. The jet flow direction was determined by sail fullness. Therefore, under the effect of the reverse thrust of the jet flow, the drag coefficient of the DER decreased with increasing sail fullness. (3) Owing to the effect of the main tail vortex downstream of the canopy and the jet flow at the crescent slots of the sail, the UER and DER featured optimal fullness values for stability. (4) Considering the deceleration and stability performances of ringsail parachutes with different sail fullness values, the aerodynamic performance of ringsail parachutes was optimal at a sail fullness value of K=1.10.[Conclusions] Through the CFD-based numerical calculations of the steady descent states of ringsail parachutes, this study explores the effects of different exhaust directions and sail fullness on the aerodynamic performance of ringsail parachutes. Our research can provide a certain reference for the structural designs and performance analyses of ringsail parachutes.

[Objective] Parachutes are widely used in aviation, aerospace, and weapon fields as an efficient and economical aerodynamic deceleration device. The drag force of a parachute mainly comes from the pressure differential on both sides of the permeable canopy. The essence of canopy permeability is that the air flows through the fabric pores to form a jet, thus affecting the flow field around the parachute and, subsequently, the aerodynamic performance of the parachute. To study the aerodynamic performance of a parachute, the micropore jet and permeability characteristics of its canopy fabric must be thoroughly investigated.[Methods] The micromodels of fabrics with high and low porosities were established on the basis of TexGen, numerical calculation of the pore jet flow under different pressure differentials was performed using computational fluid dynamics, and the numerical permeability values were compared with the experimental values. Then, the pressure and velocity of the jet domain were analyzed. The jet domain was divided into four regions according to various velocity and pressure characteristics along the central axis of the pores. Since the quantitative analysis of the jet domain under different pressure differentials was difficult, the relative pressure differential and relative velocity without the dimension parameters were proposed. On this basis, the jet characteristic parameters were proposed along with the application of the jet theory. The parameter change rule of different fabrics under different pressure differentials was analyzed. Moreover, the factors influencing the jet parameters were studied. Finally, the Levenberg-Marquardt optimization algorithm was used to fit the influence range of the jet domain based on the single-phase exponential decay function, and the experimental results were compared with the numerical results.[Results] The numerical results of the micropore jet flow field showed that:(1) The velocity of air increased within the pore and decreased after the outflow, while the pressure changed occur inversely. The pressure gradient was concentrated in the pore. (2) The jet flow field comprised four zones:velocity increase zone, velocity decay zone, wake decay zone, and wake transition zone. The changes in the velocity and pressure gradients along the direction of air flow primarily occurred in the velocity increase and velocity decay zones. The maximum velocity value of the central axis and the minimum pressure value were located in the adjacent pore throat. The flow characteristic parameters in the wake decay zone were not affected by the influence of the pressure differential. (3) When the pressure differential exceeded 200 Pa, the flow characteristic parameters in the fabric pore and the jet domain were determined only by the fabric structure. (4) The influence range of the jet domain increased with the porosity and shares an exponential decay relationship with the air permeability.[Conclusions] In this paper, the variation law of velocity and pressure in the fabric microporous jet flow domain is studied based on the numerical results of the pore jet flow field under different pressure differentials. The jet domain calculation model suitable for the parachute fabric is established. The research method proposed in this paper is highly significant in exploring the fine-flow field structure of the permeable parachute and improving the accuracy of the flow field model of the permeable canopy.

Parachutes provide rapid vertical deceleration, but are more useful if they also provide stable, controllable lateral movement. A simple, effective method for lateral motion is to cut asymmetrical slits or holes in the canopy. This study simulated flat circular parachute designs with asymmetric holes and slits. The flow fields and the drag and lateral force coefficients of several designs were compared to determine how effectively these parachutes provide lateral motion. The parachute with a 30° annular slit starting from the bottom of the canopy provides the best drag. Further design variations shortened the length of the annular seam and gradually added radial holes. In all the designs, the U-shaped slit greatly improves both the drag and the lateral motion of the parachute and gives the best stability with changes in the angle of attack.

[Objective] A parafoil is a type of parachute that can glide. With the aid of navigation and control equipment, a parafoil can approach a target point by autonomously changing its course. This capability is a great advantage over other types of parachutes in precision aerial delivery and spacecraft recovery missions. Because of these special characteristics, a parafoil is more like an aircraft than a parachute. Therefore, its design must include structural and aerodynamic design, making parafoil design complex, particularly for a large parafoil. The design method of a large parafoil is of high research value and can substantially improve the performance of the parafoil. The design method needs to be more accurate and reliable to meet the needs of the parafoil in recovery missions. In this paper, a complete set of design methods for a large parafoil was investigated, included structural and aerodynamic design methods. A structural design method for a large parafoil was first proposed, including structural composition, parameter selection, main component design, and structural framework. By investigating the design parameters of proven large parafoils, proposed values for design parameters were given. At the same time, the influence of design parameter variation on parafoil performance was also discussed. In addition, a 300 m² parafoil was designed for a launch vehicle booster with the above method. On the basis of the structural design, this paper used the numerical simulation results of an airfoil to modify the aerodynamic design method of a parafoil. The modified method can obtain the stall angle of attack of a parafoil system and the imbalance of the parafoil system with a small rigging angle before the stall, which were conducived to selecting the rigging angle in the design. A wrong rigging angle will result in a parafoil system that cannot glide, which means it is a failed design. The modified method can also obtain more accurate parafoil aerodynamic data with a change in the attack angle at various rigging angles. According to this method, the aerodynamic data of the 300 m² parafoil was acquired, and its rigging angle was determined to be 4°, which allowed for good aerodynamic performance and balance performance of the large parafoil. The verification results of an airdrop test and flight test for the 300 m² parafoil were given. Comparing the aerodynamic data in the design and the test showed that:1) The data obtained by the modified aerodynamic design method agreed well with the data in the test. 2) The parameter selection in the design, such as the rigging angle, was reasonable and feasible. 3) The structural framework of the large parafoil was sufficiently strong. The design method of the large parafoil proposed in this paper is accurate and reliable. The designed large parafoil passes the airdrop and flight tests, approving that the method can be applied to large parafoils.

[Objective] Parachutes are aerodynamic devices widely used in the deceleration and landing stages of a spacecraft.The opening process is the most critical working scenario of a parachute, wherein structural damage often occurs due to a significantly high aerodynamic load. Generally, the peak aerodynamic load during inflation and deployment of a parachute can reach more than 1.5 times the load at its steady state.Because of the high elasticity and damping characteristics of the flexible fabric materials comprising the parachute, the peak dynamic load can be effectively reduced. Thus, it is extremely important to accurately predict the mechanical characteristics of fabric materials during the opening process. In this paper, a numerical simulation method is used to study the mechanical characteristics of the fabric material in the inflation and deployment of a conical ribbon parachute. The opening process of a parachute involves a strong coupling effect of the nonlinear flexible structure and flow field. To investigate the influence of latitudinal reinforcing bands with different elasticities on the mechanics of a ribbon parachute, the opening processes of parachutes with different elastic and without latitudinal bands are simulated using the fluid-solid interaction (FSI) method. In the three parachutes, nylon and aramid fiber Ⅲ are used for latitudinal reinforcing bands in the first and second parachutes, respectively, the third has no latitudinal reinforcing band. The arbitrary lagrange-euler (ALE) method is applied to simulate the opening process, and the penalty-function method is used to demonstrate the force and displacement information between the canopy and flow field elements. The numerical simulation process is performed based on the LS-DYNA solver with the single-machine-distributed parallel computing strategy. Based on the simulation results of aerodynamics, latitudinal reinforcing band tension, and parachute canopies stress, the effect of the fabric material elasticity on the dynamic load of a parachute during the opening process is analyzed. Eventually, the aerodynamics of the parachute without latitudinal bands during the opening process is tested using the wind tunnel test, and the feasibility of predicting the mechanics of a parachute by the FSI method is verified. The simulation and experimental results showed that the elastic modulus of latitudinal bands had a nominal effect on the overall aerodynamics but had a significant effect on the stress of the canopies and latitudinal bands during the opening process. During the opening of a parachute, the maximum stress in the canopy appearaled soon after the reefing stage ends, and the projected area and aerodynamic load of the canopy increased exponentially at this time. Compared with aramid fiber Ⅲ and without latitudinal bands, the maximum stress of the canopy with a nylon configuration was reduced by 22.8% and 11.5%, respectively. Additionally, the maximum stress of the latitudinal bands and canopies made of nylon were reduced by 83.3% and 22.8%, respectively, compared to those made of aramid fiber Ⅲ. Based on the finite element method, three dynamic models of conical ribbon parachutes with different latitudinal band configurations are established. Numerical analysis of the opening process is performed, and the part-delete method is introduced to simulate the disreefing process.The parachute opening process can be effectively simulated using the FSI method, which is later applied in the selection of materials and optimization of the parachute design.

[Objective] In the practical applications of a parachute, its force state is complicated and exists in different phases, including the deployment, inflation, and steady-state descent phases. Current parachute design and verification methods based on traditional static tensioning do not match the real usage of parachute materials. Material performance verification tests for parachutes in real environments have become a major concern for researchers. To simulate the real working condition of a parachute as closely as possible in the ground material test and improve the design verification method, a test program was designed according to actual working conditions from three aspects:fatigue load, plane bidirectional load, and vertical plane load on the textile material for a parachute. Obvious breathing and surge phenomena are observed in the parachute inflation process, making the parachute fabric material experience repeated loading and unloading. To study the influence of fatigue load on the parachute strength, the fatigue test was designed for the fabric material for a parachute. In the fatigue test, two typical fabric structures were tested in 150, 250, 500, and 1 000 cycles. The parachute was in a typical multidirectional stress state during the use, whereas the existing traditional static tensile test is a one-way test. The breaking strength of the material under biaxial tensile load was tested and compared with the strength of the material under uniaxial tensile. The parachute fabric would inherit the vertical force on its surface during use, the bursting test was conducted to study the fabric's ability to bear the vertical load, and the load calculation method for the fabric subjected to a vertical plane load was proposed and compared with the fracture strength of the fabric subjected to the plane load. The fatigue test results showed that the breaking strength of the fabric did not substantially change following the fatigue load, but the load reduced the elongation of nylon; 2 fabric materials were tested, and the elongation at the break of two kinds of nylon seams was reduced from about 15% to about 11% after 1 000 cycles of fatigue load, and the reduction of elongation at the break reduced the tensile breaking work. The test results of breaking strength of the material under biaxial tensile load showed no significant difference in the uniaxial tensile strength between the two kinds of nylon plain weave fabrics commonly used for parachutes. The bursting test results showed that the breaking strength of the nylon fabric under a vertical plane load was less than the nominal breaking strength of the material, with a maximum strength loss of about 16%, and the test results could be applied to the calculation method of parachute strength design factor to guide parachute design. The test results showed that under a vertical plane load, the breaking of nylon fabric in a T-shape indicates that it has a good warp and weft strength uniformity; if the breaking occurs in a zigzag shape, there is a difference in the warp and weft strength of the nylon fabric. In the actual use environment of a parachute, the fabric is subjected to the coupling effects of vertical plane load and fatigue conditions; thus, the changes in the breaking strength and elongation at the break must be considered comprehensively.

[Objective] China's new-generation manned spacecraft test ship uses two deceleration parachutes and three main parachutes for pneumatic deceleration recovery. The main parachute bag is separated and pulled out after the deceleration parachute has completed its work. The main parachute bag is then pulled out from the parachute cabin as a key and important link in parachute deceleration in the design of the recovery landing system. The pullout process of the main parachute bag involves multibody coupling, such as slings, parachute bags, and hatch covers. Thus, the process is relatively complicated and theoretical calculations are generally used. Accurately describing the process, which could lead to countless errors in the actual process, is difficult.[Methods] This paper propose a finite element analysis method to establish a dynamic analysis model for the main parachute out of the cabin to solve the aforementioned problem and accurately understand the changes produced in this process. In addition, the actual parachute and rope system calculation model is established to quantify the relative motion and load of each component accurately in the pullout process of the main parachute, and the load distribution of different connecting parts is obtained, providing support for the system design and evaluation. The parachute aerodynamic deceleration model is simplified on the basis of the aerodynamic load dynamic matching control method, which affects the initial speed of the main parachute out of the cabin. Influencing factors, such as the tensile force of the heat shield and the quality of the hatch cover, are comprehensively analyzed and compared. The response characteristics of the load, speed, and overload of the main parachute bag out of the cabin are obtained, and the entire process of the main parachute bag out of the cabin is intuitively described.[Results] The method for the main parachute out of the cabin demonstrated the following results:1) The main parachute pullout process based on finite element analysis technology could accurately describe the coupling relationship between parachute components and simplify the aerodynamic calculation of parachutes. The dynamic iterative aerodynamic load subroutine could reduce the calculation amount and provide an efficient analysis means for the dynamic analysis of the reentry capsule. 2) With the increase in the initial speed in the pullout process of the main parachute, the time required for the complete pullout of the main parachute was short, the load on each sling increased, and the overload of the entire cabin also showd a rising trend. Therefore, the initial drawing boundary conditions should be strictly controlled. 3) The mass of the main umbrella covered considerably affected the peak force of the total towing load curve of the umbrella and would increase with the mass. Thus, the weight of the main umbrella cover should be strictly controlled in the scheme design to reduce the load of the umbrella towing.[Conclusions] This method effectively guides the design of the recovery system for a new generation of manned spacecraft test ships and provides theoretical support for subsequent formal flight missions.

[Objective] The conical deceleration structure is a typical shape in inflatable reentry and descent technology (IRDT). Compared with the traditional rigid deceleration structure, the inflatable deceleration structure represented by alumina fiber has lower heat resistance. Therefore, accurate thermal environment prediction is crucial for designing the IRDT system. Moreover, high pressure deforms the surface of the inflatable structure, so the surface pressure distribution is another issue that needs attention. The surface heat flux and pressure distribution of a conical deceleration structure under thermochemical reaction conditions are analyzed through numerical simulation. At the same time, the influence of different half-cone angles on the surface heat flow and pressure distribution is analyzed.[Methods] The numerical model is based on the integral Navier-Stokes (N-S) equation. The Park85 and the two-temperature nonequilibrium models are used to calculate the thermochemical reaction with a noncatalytic wall condition. The equations are solved using the finite volume method. The lower-upper symmetric Gauss-Seidel method is adopted for iteration. The blunt body standard model ELECTRE is used to validate the numerical model. The calculation case of a conical deceleration structure with a height of 70 km is investigated, and the inlet Mach number is 13. The variations in temperature and chemical component concentration along the stagnation line, as well as heat flow and pressure distributions on the structure surface are studied. In addition, the simulation of four conical deceleration structures with different half-cone angles is carried out to analyze the effect of the half-cone angle on the surface heat flow and pressure.[Results] The simulation results show that 1) the gas translational temperature after the shock wave is approximately 7000 K. Along the stagnation point line, the vibrational temperature gradually increases, and the two temperatures reach equilibrium near the stagnation point and decrease to the wall temperature. 2) The concentration of the N component in the shock layer is low and decreases to 0 at the stagnation point. The O and NO components gradually increase along the stagnation point line and reach the maximum near the stagnation point. 3) The surface heat flow and pressure are the highest at the stagnation point and decrease rapidly along the radial direction near the stagnation point. Then, the heat flow decreases linearly, and the pressure is approximately constant. 4) For different half-cone angle conical deceleration structures, the shock wave positions and surface heat flow distributions of the 50°, 55°, and 60° cases are basically identical. The shock wave position of the 65° case is farther from the leading edge, and the surface heat flux is lower. 5) Finally, the stagnation pressures of the four cases are basically identical, and the peripheral pressures increase linearly with increasing half-cone angle.[Conclusions] The surface heat flow and pressure distributions on the conical deceleration structure can be revealed by the numerical calculation. The change in the half-cone angle significantly impacts the surface heat flow and pressure distributions of the conical deceleration structure.

[Objective] As an emerging space return technology, the inflatable reentry vehicle provides a new technical approach to deep space exploration, recovery of target spacecraft, and freight transport of space products. The inflatable reentry vehicle faces severe aerodynamic load impact, and many structural safety failures have been induced by the aeroelastic effect in actual test flights. However, the existing structural finite element models for inflatable reentry vehicles do not fully consider the inflation gas mechanism and material nonlinearity and cannot accurately describe the nonlinear structural dynamic characteristics of flexible inflatable structures. The defect of the low fluid-solid coupling degree also exists in the aeroelastic system, and how the inflation gas participates in the process of coupling is unclear. The limitations of existing methods impede reasonably revealing the aeroelastic characteristics under the influence of external high-speed flow and internal inflation gas during reentry.[Methods] Aimed at the aeroelastic dynamic response of the inflatable space reentry aeroshell in supersonic and transonic flows, a fluid-solid coupling model considering inflation gas is established in this paper, which also considers the influence of structural deformation on the flow field more than existing methods, and the LES can effectively describe the flow field with strong separation characteristics. Meanwhile, the six-DOF flight dynamics are used to modify the flight trajectory in the supersonic stage, and the two-way coupling between flight dynamics and aerodynamics is effectively realized. The proposed method can reveal the dynamic response characteristics under the action of large aerodynamic force and inflation gas, which is closer to the physical essence of aeroelasticity.[Results] The results indicate that the vehicle will vibrate violently in the transonic and supersonic flow fields, which is essentially the buffeting effect under the action of large-scale turbulent wake vortices. Under a Ma 0.8 flight condition with the most severe airflow pulsation, the axial and pitching vibration amplitudes of the vehicle reach 40 mm and 67 mm, respectively. The frequencies of airflow pulsation and structural vibration are relatively low, leading to a potential risk of inducing resonance with the natural frequency of the vehicle. Under transonic and supersonic conditions, the aerodynamic moment derivative of the vehicle is negative when the attack angle is less than 50°, and the structure can maintain static stability.[Conclusions] According to the calculation results of the aeroelastic dynamic response, the amplitude of the structure also converges after the release under an attack angle of 17°, which further confirms that the aerodynamic instability will not be severe. In the transonic and supersonic regions, with decreasing Mach number, the Reynolds number and inertia force increase continuously. Because of the increase in inertia force, the separation point at the shoulder moves forward, and the position of the shear layer moves outward, which increases the wake width and enhances the vibrational amplitude of the structure. When the attack angle exists, the increase in flow mixing due to asymmetric flow in the upper and lower half regions is the main reason for the increase in the unsteady degree of wake and the vibrational amplitude. This research provides a valuable reference for inflatable space reentry aeroshell structure safety design and evaluation under transonic and supersonic flows.

[Objective] The Tianwen lander has to follow the same sequence as most space missions landing on other planets (or back on Earth), a process known as Entry, Descent, and Landing. NASA engineers have described the descent of the Mars landing missions as "seven minutes of terror" as it is the most unpredictable. About 20 attempts to land on Mars have been made by different countries so far. Besides a considerably thinner atmosphere, Mars's gravitational field is weaker than that of Earth; thus, on average, it delivers 38% as much downward acceleration. In this paper, a large-scale gravity compensation system is designed to simulate the Martian environment to test the Tianwen-1 lander during the Mars landing.[Methods] Considering severe collisions and abrupt changes of states during the landing, the system uses multiple elastic elements, including springs, to eliminate undesired high-frequency vibrations, thereby enabling the system to maintain a stable output during severe dynamic processes. The compensation system is composed of an adjustment mechanism, springs, guide bar, wire rope, and oscillating bar. To achieve the target stiffness, five springs are used in parallel. By adjusting the adjustment mechanism, the initial preload of springs can be varied to match different loads with masses varying within a certain range. The guide bar can restrain the lateral movement of the spring, thereby ensuring that it maintains a stable state during shortening and elongation. Additionally, it can also offset the weight of the springs. Conversely, the equivalent replacement of the zero-free-length spring enlarges the stroke of the system. To achieve the equivalent replacement of the zero-free-length spring, an additional mechanism and pulley are designed. Then, the mechanical properties are explored from the perspective of energy conservation. Eventually, the relationships among the characteristic values of each component in the system can be determined.[Results] Three major issues had been resolved by the gravity compensation system. 1) With the lander moving at high speed, the system successfully achieved gravity compensation for heavy loads (7 200 N) in a long stroke (800 mm). The error between the experimental and simulation results was within the allowable range. 2) When the lander hit the ground, the system output a constant force (maximum error:7.8%), thereby implying that the system had good adaptability for dynamic processes. 3) During the entire landing process, the tracing accuracy (maximum error:7.8% and average error:1.5%) of the constant-force output from the system had already met the requirements (maximum error:10% and average error:10%).[Conclusions] To fully or partially compensate for the gravity of landers during the landing process, this paper presents a large-scale gravity compensation system. The design, analysis, fabrication, and experimental testing are implemented to investigate the performance in terms of the constant-force output. With the Mars lander descending at high speed, the system successfully achieved gravity compensation for the heavy load (7 200 N). During the landing process, the tracking accuracy of the output force of the system has already met the requirements. Furthermore, the compensation system can be quickly adjusted to suit a target planet.

[Objective] Recently, a trend has been developed toward the high-density launches of launch rockets, and unprecedented attention has been paid to the impact zone safety of rocket's separated fairings. Great pressure on controlling the environmental safety of the rocket's fairing half results from the low accuracy and large uncertainty in reentry trajectory calculation using the traditional mass point ballistic model. The main reason for the difficulty in reliably calculating the reentry ballistics of the fairing half is the lack of comprehensive studies on the half fairing's aerodynamic characteristics under various reentry flight conditions. This study aims to deal with the problem of reliable prediction of the reentry impact point of fairing half after separation from the launch vehicle. By using the computational fluid dynamics (CFD) approach, a comprehensive study is conducted here on the aerodynamics of the fairing half of a common rocket in the continuum regime of its reentry process. The incoming flow parameters for multiple computational conditions are extracted from the measured ballistic data of a certain flight mission. The calculation of various conditions is set and controlled by a parametric automated script. To obtain the relevant aerodynamic coefficients, the steady-state numerical solution for three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations is obtained using the finite volume method. The Roe scheme and the implicite lower-upper symmetric Gauss-Seidel (LU-SGS) solution algorithms are used to obtain the discretization solution of the flow control equations. A numerical model uses an unstructured mesh structure, with a total mesh number approximating 15 million, and generates enough prismatic layers on the surface. The turbulence model is a two-equation realizable k-ε model. Aerodynamic coefficients of the fairing half were obtained under various flight conditions, where the Mach number varied from 0.20 to 5.95, while the angle of attack (AOA) varied from 0° to 360°. Numerical results indicated that:1) Two trim angles of attack existed in the supersonic and hypersonic regions, with the first trim AOA ranging from 85° to 97° and the second trim AOA ranging from 256° to 254°. 2) Similarly, two trim angles of attack existed in the transonic and subsonic regions, with the first trim AOA ranging from 85° to 88° and the second trim AOA ranging from 259° to 252°. 3) Whether in the supersonic or transonic region, the fairing half behaved statically stabled at its first trim AOA in axial roll direction at 0° roll angle but non-statically stabled at the second trim AOA. 4) Adjusting the position of the center of mass of the fairing half along its axial direction could effectively change its trim AOA, which led to a significant change in the lift-to-drag ratio in turn. However, adjusting the position of the center of mass along its radial direction had little effect on the trim AOA and the lift and drag characteristics. By utilizing the obtained database of aerodynamic coefficients, the 6-degrees-of-freedom reentry model for the fairing half can be achieved, helping improve the prediction accuracy of the impact area significantly. In the hypersonic region, the difference in the aerodynamic coefficients under the same AOA condition is below 15%, indicating that in the continuum region, when the Mach number is greater than 5.95, the required aerodynamic coefficients for the ballistic reentry analysis are similar to those of Mach number is 5.95. Due to the obvious differences in the aerodynamic coefficients in the transonic or subsonic region, the required aerodynamic coefficients for the ballistic reentry analysis can be interpolated from the obtained database. The trim AOA and the corresponding lift-to-drag ratio can be effectively changed by adjusting the position of the center of mass of the fairing half along its axial direction, thus adjusting or controlling the impact point within a certain range.

[Objective] The hybrid unmanned aerial vehicle (UAV) has become an important technical means in the UAV field because of its excellent lift-drag characteristics and endurance. The endurance of an UAV can be improved mainly by obtaining good lift and drag characteristics and reducing its weight. Using a modular UAV as the research object, this paper established a hybrid UAV configuration concept, mainly considering the stress change rule and the maximum fatigue life of the monomer UAV under aerodynamic force, and its lightweight design goals under the fatigue life and the maximum stress constraint. This paper conducted research on lightweight design mainly in three ways:1) According to the hybrid UAV concept, an overall configuration of "main body + monomer" is proposed and the full-text research object is provided. The monomer UAV is an important task-combined UAV unit, so the mechanical properties of every single part and the overall topology structure are emphasized. 2) In the incompressible unsteady three-dimensional continuous equation and Navier-Stokes (N-S) equation, momentum analysis is carried out on the aerodynamic characteristics of a single wing UAV in a full working environment, the regional flight parameters are clarified, and the maximum aerodynamic force is selected as the ultimate load, in the form of loading a sine function into the single wing. The monomer stress distribution of the maximum stress region of the wing-body is extracted. Combined with Fe-safe numerical analysis software, fatigue life analysis is performed, and improvements are proposed for sensitive areas. Based on the service life guarantee and the solid isotropic microstructure with the penalization (SIMP) method, a topological optimization analysis of a single wing, load-bearing frame is performed. 3) A monomer wing finite element model is established to verify the correctness of the topology optimization model. For a single wing, all the structural analyses of the static and dynamic analysis results verify the correctness of the above theoretical analysis. At the same time, in the topology analysis link, because each tolerance beam differs in weight, analysis was performed for each beam. In this paper, based on the classic SIMP topology optimization method, the topological calculation and analysis of the single wing's front beam, middle beam, rear beam, and middle rib show that the weight of each beam is reduced by 5%, 10%, 5%, and 15%, respectively, under the premise of invariable mechanical distribution. After obtaining the analysis results, the weight of the entire wing is reduced by 35%. an innovative topology optimization process is adopted to analyze the fatigue life of key parts before the wing is joined, ensuring the process optimization of the wing and the service life of key parts. The proposed method integrates key position optimization, topology optimization, and fatigue life analysis, avoiding separate post-assembly fatigue analysis of key positions, optimizing the whole analysis process, and improving the work efficiency of lightweight analysis. The wing topology optimization method under the fatigue life constraint proposed in this paper reduces the weight of a single UAV by 35% under the condition of constant force and improves the efficiency of wing lightweight analysis. The new analysis method provides technical support for subsequent UAV optimization analysis.

[Significance] Neither a global atmosphere nor a magnetic field exists above the lunar surface. The lunar surface is exposed to high-energy ultraviolet radiation and solar wind, which charge a large amount of lunar dust to form high-potential, local electric fields. The lunar mountains, with a length of more than ten kilometers and a vertical height of several kilometers, block ultraviolet radiation from the sun, creating a whole light-dark junction region with a shadow region on the lunar surface. Meanwhile, the occlusion of the lunar surface by a detector or a rover also forms a local light/dark junction region under sunlight. Under the lunar environment, complex electric fields exist at the whole and local light-dark junction regions, which leads to the abnormal phenomenon of charged lunar dust transport. In this paper, the whole and local light-dark interface models were proposed, and the particle-in-cell method and leap-frog method were used to study the lunar dust transport anomalies of the two models based on measured data of the Apollo lunar dust samples. The results showed that a very significant abnormal phenomenon of lunar dust transport occurred in the whole and local light-dark interface models. First, the lunar dust transport path was bell-shaped or parabolic, and a large amount of lunar dust accumulated above the boundary in the whole light-dark interface model. Moreover, the lunar dust transport phenomenon had an obvious horizontal transport characteristic, and the horizontal transport velocity was approximately two to ten times that of the vertical direction. Particularly, the abnormal phenomenon of lunar dust transport at the junction between light and dark was more severe and more likely to occur in the whole light-dark interface model. Second, in the local light/dark interface model, significant anomalies of lunar dust transport were found on both the left and right side of the detector, where a large amount of lunar dust was transported bidirectionally from one side to another side, passing through the detector in the horizontal direction. In particular, a large or small lunar dust vortex rotated around both sides of the left and right detector, resulting in a local "moondust storm" around the detector with a number density of approximately 1.4×10⁵ particles per cubic meter. This paper predicts the abnormal phenomenon of lunar dust transport for the whole light-dark junction region caused by mountain occlusion, which can indirectly verify the observed occurrence of the "horizontal glow" phenomenon. Abnormal transport of lunar dust in the local light-dark junction region may be one of the main reasons for the deposition of much lunar dust on the detector/rover and so on. The abnormal phenomenon of lunar dust transport is not only potentially harmful to the current lunar detector/rover and so on, but also a key problem that human exploration activities cannot avoid. The results of this study have not only an important reference value for selecting the landing site of a lunar detector and the traveling route of a rover but also helpful to reduce the lunar dust pollution caused by the detector/rover and so on.

[Objective] Lunar and Mars exploration projects face the issue of autonomous spacecraft landing or taking off on the surface of the target celestial body in a low-gravity environment. Several mechanical problems arise when a spacecraft is flying in a low-gravity environment, which is different from Earth's gravity environment. To solve these problems, it is necessary to verify the key actions of spacecraft for soft landing on or taking off from the target celestial body.[Methods] Since the Chang'e-3 mission of the second stage of lunar exploration, the test system of the suspension low-gravity simulation technology has been studied via fast follow-up and multilevel compensation, and the real-time tracking technology for the spacecraft motion in large space has been developed to keep the pulling force applied in the entire process constant and the deflection angle of the force direction relative to the vertical direction sufficiently small. Through the technical improvement of the third stage of the lunar exploration mission and the technical upgrade iteration of the Mars exploration, a complete set of technical systems for the spacecraft has been successfully constructed to simulate the low-gravity landing and take-off test of extraterrestrial celestial bodies on the earth. A three-dimensional (3D) follow-up system adopts the two-level linkage driving technology of a large-scale follow-up and rapid and accurate tracking to construct a landing and take-off test method for simulating the low-gravity environment of a spacecraft on the ground, thus overcoming a series of technical difficulties, such as large test space and high control accuracy, and employing several key technologies, including multi degree-of-freedom linkage of the 3D follow-up system and high-speed and high-precision coordinated control of the large-inertia electromechanical equipment.[Results] The large-scale follow-up tracking of the spacecraft in the test process was achieved by controlling the movement of the rapid follow-up platform through the parallel-link system driving technology, and the requirement of the absolute inclination angle of the lifting rope was realized by applying a high-precision tension control to the spacecraft through the rapid follow-up platform device and following the movement of the spacecraft in the horizontal direction. Additionally, the horizontal stiffness of the fast follow-up platform was improved to overcome the adverse effects of the coupling shaking of the two-level linkage equipment in the spacecraft test.[Conclusions] The system has been successfully applied to a series of real ground test conditions, such as hovering, obstacle avoidance, slow descent, landing, and take-off of China's Chang'e-3 and Chang'e-5 in the lunar exploration project and Tianwen-1 spacecraft in Mars exploration. The test data which can support the research and engineering exploration of spacecraft are obtained, providing a key technical means for verifying and optimizing the comprehensive performance parameters of the spacecraft. With the continuous development of space missions, the mechanical environment simulation and ground test technology regarding spacecraft landing and take-off from extraterrestrial bodies pose new challenges. The 3D follow-up system for a low-gravity simulation will further develop toward a high-precision, large-load, and high-dynamic simulation technology, laying the foundation for the application of ground low-gravity simulation tests for manned lunar landing and deep space exploration missions.