Simulation of hollow shaped charge jet impacts onto aluminium whipple bumpers at 11 km/s

The computational technique of Smoothed Particle Hydrodynamics (as implemented in the hydrocodes AUTODYN-2D and AUTODYN-3D) has been used to simulate the impact of hollow shaped charge jet projectiles onto stuffed Whipple bumper shielding. Due to limited availability of material models, the interim Nextel/Kevlar-Epoxy bumper was modelled as an equivalent thickness of aluminium. Stuffed Whipple bumper shields are used for meteoroid and debris impact protection of the European module of the International Space Station (the Columbus APM). A total of 56 simulations were carried out to investigate the impact processes occurring for shaped charge jet impact. Sensitivity studies were carried out on the influence of projectile shape, pitch, yaw and strength at 11 km/s to determine the range of debris cloud morphologies. The debris cloud structure was shown to be highly dispersed, and no projectile remnant was observed at the centre of the cloud. The mass of an aluminium sphere producing equivalent damage to a shaped charge jet projectile was in the range 1.5 to 1.75 times greater than the mass of the shaped charge jet projectile. Upon loading by the dispersed debris cloud, the interim bumper failed by spallation, producing fragments moving at 2 km/s or less. The fragments distorted the rear wall (pressure wall) of the shield but did not perforate it. The experimental data show rear wall deformation but to a lesser degree. Perforation of the rear wall, observed for one test, was not reproduced by the simulation. Nextel/Kevlar-epoxy material models are required to reproduce correctly the interim bumper failure under debris cloud loading.     

Three dimensional hypervelocity impact simulation for orbital debris shield design

This paper assesses a Whipple shield impact simulation method which is both accurate and computationally efficient. The paper documents the simulation methodology and results of Whipple shield simulations at an oblique impact angle of 30°. These results are compared with HVI experiments to demonstrate the accuracy of the simulation technique. In addition, simulations of Whipple shields in the velocity regime above 8km/s were completed and the results compared to published ballistic limit equations to demonstrate the reliability of these equations. Finally, the paper documents computational efficiency of the simulation technique.     

泡沫铝防护屏的Whipple防护结构弹道极限数值模拟研究

为研究泡沫铝板作为防护屏的Whipple防护结构抵御空间碎片超高速撞击的特性,模仿泡沫金属的生产原理建立了泡沫金属细观结构几何模型,结合自编的光滑质点流体动力学程序进行了超高速撞击数值仿真,通过与实验对比验证了模型的有效性.分别对相对密度为23.2%的理想均匀和非均匀开孔泡沫铝板作为防护屏的Whipple防护结构进行了数值仿真,得到了它们的弹道极限曲线,并与实心铝板作为防护屏的Whipple防护结构进行了对比分析.结果表明,相同面密度的泡沫铝板相对于实心铝板能够在更低的速度上将弹丸粉碎、液化及气化.泡沫铝板作为防护屏,在总体上拥有更好的防护性能;相同面密度的理想均匀泡沫铝板的防护性能总体上优于非均匀泡沫铝板.     

Performance of Whipple shields at impact velocities above 9 km/s

The results of 18 impact tests performed on Whipple shields were compared to the predicted ballistic limits of the shields in the region where the impact velocity of the threatening particle was high enough to produce melting and incipient vaporization of the particle. Ballistic limit equations developed at NASA Johnson Space Center were used to determine nominal failure thresholds for two configurations of all-aluminum Whipple shields. In the tests, 2017-T4 aluminum spheres with diameters ranging from 1.40 to 6.35 mm were used to impact the shields at impact velocities ranging from 6.94 to 9.89 km/s. Two different aluminum alloys were used for the rear walls of a simple Whipple shield. The results of 13 tests using these simple Whipple shields showed they offered better-than-predicted capability as impact velocity increased and that the strength of the rear wall material appeared to have a smaller-than-predicted effect on the shield performance. The results of five tests using three configurations of a scaled Space Station shield - a plain shield at 0 degrees, two shields with multilayer insulation in the space between the bumper and the rear wall (also at 0 degrees), and two tests with the plain shield at 45 degrees obliquity - showed that these shields met their predicted capabilities.     

Space station JEM design implementation and testing for orbital debris protection

The Japanese Experiment Module (JEM) is the Japanese contribution to the International Space Station (ISS) Program. The core part of JEM is a Pressurized Module where the crew conducts space experiments in a microgravity environment in space. The development of a shield design to protect against micrometeoroids and orbital debris (MM/OD) has been a key issue for the permanent manned space station which is expected to have an operational life exceeding 10 years. Many technical approaches for MM/OD protection have been studied. As the launch of the space station elements draws near, the shield design has become mature, and the technical test data for the MM/OD shielding obtained has increased confidence in its performance. NASDA, which is responsible for JEM design and development, has conducted a series of tests and simulations to define the MM/OD shield configuration and design. The structural failure due to MM/OD strikes has also been assessed for the combined shield and pressure shell configuration.     

Hypervelocity impact tests and simulations of single whipple bumper shield concepts at 10km/s

A series of experiments has been performed to evaluate the effectiveness of a Whipple bumper shield to orbital space debris at impact velocities of 10 km/s. Upon impact by a 19 mm (0.87 mm thick, L/D 0.5) flier plate, the thin aluminum bumper shield disintegrates into a debris cloud. The debris cloud front propagates axially at velocities of 14 km/s and expands radially at a velocity of 7 km/s. Subsequent loading by the debris on a 3.2 mm thick aluminum substructure placed 114 mm from the bumper penetrates the substructure completely. However, when the diameter of the flier plate is reduced to 12.7 mm, the substructure, although damaged is not perforated. Numerical simulations performed using the multi-dimensional hydrodynamics code CTH also predict complete perforation of the substructure by the subsequent debris cloud for the larger flier plate. The numerical simulation for a 12.7 mm flier plate, however, shows a strong dependence on assumed impact geometry, i. e., a spherical projectile impact geometry does not result in perforation of the substructure by the debris cloud, while the flat plate impact geometry results in perforation.     

A combined experimental and numerical approach to study ballistic impact response of S2-glass fiber/toughened epoxy composite beams

A combined experimental and 3D dynamic nonlinear finite element (FE) approach was adopted to study damage in composite beams subject to ballistic impact using a high-speed gas gun. The time-histories of dynamic strains induced during impact were recorded using strain gages mounted on the front of the composite beam specimen. During ballistic impact tests, the impact velocity was also measured. The commercially available 3D dynamic nonlinear FE code, LS-DYNA, modified with a proposed user-defined nonlinear-orthotropic damage model, was then used to simulate the experimental results. In addition, LS-DYNA with the Chang–Chang linear-orthotropic damage model was also used for comparison. Good agreement between experimental and FE results was found from the comparisons of dynamic strain and damage patterns. Once the proposed nonlinear-orthotropic damage model was verified by experimental results, further FE simulations were conducted to predict the ballistic limit velocity (V₅₀) using either the number of damaged layer approach or a numerically established relation between the projectile impact velocity versus residual velocity or energy similar to the classical Lambert–Jonas equation for metals.     

Shielding against space debris. A comparison between different shields: The effect of materials on their performances

The space environment requires the Space Station to be shielded against orbital debris. A technological programme undertaken by the European Space Agency has led to a preliminary definition of the shield configuration for the European Attached Pressurized Module. The envisaged shield is a modified Whipple shield. A second bumper is located midway between the first bumper and the backwall.

The work described has been initiated to quantify experimentally the merits of different shields compatible with the APM system requirements. For this technological investigation, two requirements had to be satisfied. The spacing between the front bumper and the backwall had to be limited to 120 mm. The backwall thickness could not be reduced to technological limits as it has structural functions as well. In addition, the long life requirements of the Space Station precludes the use of unproved materials for the external parts of the shield.

Different materials have been tried as second bumper. The effect of the first bumper thickness on the projectile fragmentation has been explored as well. Shields based on Aluminium, Kevlar and Glare have been investigated. Kevlar 29 fabrics impregnated with epoxy resin were used for this work. Glare is a material developed to improve the fatigue strength of metal structures. It is primarily intended for aircraft skin applications. Glare consists of a 60 percent fibre volume adhesive prepreg with high-strength unidirectional or cross-ply R-glass fibres. A variety of lay-up sequences is available ranging from 2/1 (two layers of aluminium alloy sheet bonded by one layer of prepeg) to any number of layers. The 2/1 layers version of the Glare material has been used for this work.

The tests results indicate the performances of materials can change significantly with the impact conditions. Glare shows the best performances in the low velocity regime while Kevlar is very promising in the high velocity regime. It is concluded the use of Kevlar can improve substantially the performances of the APM shield.     

Ballistic limit evaluation of advanced shielding using numerical simulations

The advanced shielding concept employed for the Columbus module of the International Space Station consists of an aluminum bumper and an intermediate shield of Nextel and Kevlar-epoxy. Until recently, the lack of adequate material models for the Nextel cloth and Kevlar-epoxy has precluded the practical usage of hydrocodes in evaluating the response of these shields to hypervelocity impact threats. Recently hydrocode material models for these materials have been proposed [1,2] and the further development and completion of this model development is reported in this paper. The resulting models, now implemented in AUTODYN-2D and AUTODYN-3D, enables the coupling of orthotropic constitutive behavior with a non-linear (shock) equation of state. The model has been compared with light gas gun tests for aluminum spheres on the advanced shield at impact velocities between 3.0 and 6.5km/s [3]. Reasonable correspondence has been obtained at these impact velocities and thus the models have been used to perform preliminary assessment of predicted ballistic limits at velocities from 7 to 11km/s. The predicted ballistic limits are compared with ballistic limit curves derived on the basis that damage is proportional to projectile momentum     

Development of a lightweight space debris shield using high strength fibers

In order to protect space structure against space debris impacts, it is indispensable to develop a shield with high strength materials. A high strength fiber is one of potential materials from a viewpoint of strength, lightweight, and flexibility. The purpose of this study was to develop a new lightweight shield composed of high strength fibers against medium size debris impacts. We developed four kinds of shields using Vectran fibers, and hypervelocity impact tests were carried out by a railgun accelerator. The experimental results showed that the developed shield could stop the polycarbonate projectile with 13 mm in diameter, 1 gram in weight, and 6.9 km/sec in velocity. Adoption of the high strength fiber in the bumper materials may reinforce the protection capability and reduce the weight drastically.     

Laboratory simulation of hypervelocity debris

A series of hypervelocity damage experiments were performed on spacecraft materials in order to simulate micro-size space debris traveling at 3 to 8 km/s. Two types of impact simulations were investigated: high-power pulsed laser and laser-launched micro-flyer plate. In the first case a laser was used to generate a high-pressure shock wave which propagated into the target by means of rapid ablation of the target surface. The second case used the same laser to accelerate micro-flyer plates at a target. The laser-ablation technique and the apparatus used to propel the micro-flyer plates were compatible with a space environmental chamber equipped with instrumentation capable of analyzing the vapor ejected from the sample. Data obtained from two separate damage effects were of interest in this study: the vapor blow-off produced by the impact and the mechanical damage to the target. The value of the data obtained from both simulation methods was evaluated in terms of likeness to actual space debris damage.

Data for this work were obtained from polysulfone resin and a graphite polysulfone composite. Polysulfone was selected because it was flown on the Long Duration Exposure Facility (LDEF) satellite which spent several years in low earth orbit and experienced many space debris impacts.

The chemistry of the vapor produced by the two simulation techniques was analyzed with a time of flight mass spectrometer (TOFMS) which measured changes in the vapor chemistry as a function of time after impact, obtained a velocity measurement of the vapor, and estimated surface temperature immediately after impact using dynamic gas equations. Samples of the vapor plume were also captured and examined by transmission electron microscopy (TEM).

The mechanical damage effects caused by the simulation methods on a graphit polysulfone composite and a polysulfone resin were studied. Impact craters were examined under optical and scanning electron microscopes (SEM). Based on the two damage effect criteria the micro-flyer method proved to be a useful way to simulate hypervelocity impact of space debris. The laser-ablation method however, had shortcomings and required drastic compromises in the set criteria.     

Crater distribution on the rear wall of AL-Whipple shield by hypervelocity impacts of AL-spheres

All spacecraft in low orbit are subject to hypervelocity impact by meteoroids and space debris, which can in turn lead to significant damage and catastrophic failure. In order to simulate and study the hypervelocity impact of space debris on spacecraft through hypervelocity impact on AL-Whipple shield, a two-stage light gas gun was used to launch 2017-T4 aluminum alloy sphere projectiles. The projectile diameters ranged from 2.51 mm to 5.97 mm and impact velocities ranged from 0.69 km/s to 6.98 km/s. The modes of crater distribution on the rear wall of AL-Whipple shield by hypervelocity impact of AL-spheres in different impact velocity ranges were obtained. The characteristics of the crater distribution on the rear wall were analyzed. The forecast equations for crater distribution on the rear wall of AL-Whipple shield by normal hypervelocity impact were derived. The results show that the crater distribution on the rear wall is a circular area. As projectile diameter, impact velocity and shielding spacing increased, the area of crater distribution increased. The critical fragmentation velocity of impact projectile is an important factor affecting the characteristics of the crater distributions on the rear wall.     

A comparison of ballistic limit with adaptive-mesh Eulerian hydrocode predictions of one- and two-plate aluminum shielding protection against millimeter-sized Fe–Ni space debris

Hypervelocity collisions with space debris (SD, natural meteoroids and man-made artifacts) can significantly affect the performance of spacecraft. Here, I compare (1) the predictions of the Cour-Palais/Christiansen (C-P/C) ballistic limit equations (BLEs) spacecraft shield models with (2) the predictions of the response of those shields generated by an adaptive-mesh Eulerian hydrodynamic code, incorporating Mie-Grüneisen solid mechanics and a simple material-failure model, running on a modern PC, for hypervelocity collisions with millimeter-sized iron–nickel (Fe–Ni) spheres. The results show that the shield thicknesses predicted by the C-P/C BLEs are consistent with the adequacy of the shield response predicted by the hydrodynamic modeling. Although several hydrocodes have been used to validate the C-P/C BLEs, validating them with an (inherently computing resource-efficient) adaptive-mesh Eulerian hydrodynamic code for this impact regime appears to be novel.     

The shape effect of non-spherical projectiles in hypervelocity impacts

This paper provides qualitative and quantitative analyses of regular non-spherical projectile hypervelocity impacts on basic Whipple shields using test data obtained by light-gas guns, flat plate accelerators and shaped charge launchers. Surrogate cadmium and zinc test results are used to extend light-gas gun data beyond 8 km/s. Advanced Whipple shield derivatives are shown to be necessary to protect against non-spherical projectiles.     

Numerical simulation of non-perforating impacts on shielded gas-filled pressure vessels

In order to calibrate the output of hydrocode simulations of hypervelocity impacts on shielded gas-filled pressure vessels, Light Gas Gun impact experiments were performed. In a first step, tests were performed on so-called equivalent Whipple shield (EWS) configurations having basically the same set-up as the shielded pressure vessels (i.e. bumper thickness and - material, stand-off and backwall plate thickness and -material). Purpose was the determination of the impact conditions that lead to penetration into the backwall plate but not perforation of it or leakage through the impacted area. In a second step, impact tests on the corresponding shielded pressure vessels were performed with the same test conditions as the EWS. The purpose of the tests was the investigation whether leakage occurs when the vessel's front wall is not perforated, but just cratered. The test conditions lead to no leakage in all tests. The most important measured damage parameter was the crater depth of the deepest crater in the vessel's front wall/the backwall plate of the EWS, respectively. Hydrocode simulations were then performed to assess the capability of the numerical tool to correctly predict the damage on the impacted vessel surface. Normal impacts of aluminium spheres against shielded vessels were simulated using AUTODYN-2D, including and evaluating the effect of the static stress induced in the vessel walls by the inner pressure. Particular attention was focused on the exact determination of the maximum crater depth caused by the debris cloud impact on the vessel wall/the backwall plate of the EWS, respectively. Bumper and projectile were represented by SPH particles, the vessel shell was represented by a Lagrange grid. The results showed a very good agreement with the measured crater depths of the experiments.     

Integral model for the description of the debris cloud structure and impact

The purpose of the present paper is to introduce a new integral model capable to describe the evolution of the debris clouds originated after normal-impacts of orbital debris over a Whipple shield. This work had been developed at Alenia Spazio in the context of a degree thesis. Several numerical SPH simulations of debris impacts on a Whipple shield configuration were performed to determine the ballistic limit and to compare it with semi-empirical damage equations. In the present paper, the numerical simulations were used to investigate the typical behaviour of experimental debris clouds ([6] and [9]) and to support the development of the integral model.

With respect to previous papers ([1], [2], [3], [10]) in which a spherical shell-wise debris cloud was considered, here we try to introduce more realistic assumptions. We approximate the cloud's shape also introducing ejecta veil effects, which produce a multiplication of the deposited momentum upon the underneath wall. In the present model, the most peculiar hypothesis is a cinematic self-similar behaviour that is, whatever the shape is, the debris cloud evolves keeping unchanged its shape. Then, the material is opportunely distributed inside a volume and the choice of that distribution is described taking into account the results of the numerical simulations. Knowing the spatial material distribution and treating the cloud as a fluid, we can estimate the load time history and the drag-unitary force induced by the cloud impacting upon the rear wall. Of course, such a method uncouples the dynamic response of the rear wall from the evolution of the debris cloud. The balances of mass, momentum and energy allow three global and unknown parameters to be determined. The one-dimensional theory of impact ([10]) is used to take into account the conversion of part of the initial kinetic energy into internal thermal energy. No integration of differential equations is performed since complex propagation phenomena are taken into account through the effects they globally produce. The model still presents some free parameters related to the integral formulation. These parameters cannot be calculated through any balance condition, but they must be imposed to get a good, global reproduction of the debris cloud. The choice of these parameters is still the weak aspect of the method, and it depends on the consideration of the results obtained with more sophisticated tools, as, for instance, SPH simulations. The spatially defined load time history obtained with the debris cloud integral model can be used for further analysis on the back up plate.     

Measuring hypervelocity impact velocity from micrometeoroid crater geometry

Guided by half-space computer simulations showing hypervelocity impact crater formation for an iron particle impacting an aluminum target and characteristic crater geometry changes with impact velocity over the range 8–40 km s⁻¹, we examined normal surface crater views and cross-sectional views through craters (>0.5 mm diameter) from samples retrieved from the NASA LDEF satellite and examined in the scanning electron microscope (SEM). While geometrical features suggested in the computer simulations were indeed observed for micrometeoroid craters in 6061-T6 aluminum targets and 303 stainless steel targets, there was no consistent estimate for impact velocities in any of the experimental samples, and velocity estimates based on measuring ratios of ejecta width/crater diameter and ejecta height/crater depth as well as ejecta height/crater diameter varied from 8 to 42 km s⁻¹; over the same range simulated. These results point to the need to create reference data from actual hypervelocity impact experiments in the laboratory, and systematic observation of residual crater geometries in the SEM. These experiments also demonstrate the uncertainty in assuming a fixed impact velocity for all impact craters in space materials as well as an apparent futility in attempting to correlate impacting particle velocity with post-mortem characteristics of a given crater.     

一种改进的特征匹配算法在碎片云测量建模中的应用

为满足超高速撞击典型Whipple防护构型的损伤评估需求,利用图像处理技术对碎片云序列阴影图像进行深入研究.使用超高速序列激光阴影成像仪得到三组不同实验条件下碎片云发展过程的高清阴影图像,分别对每组最具代表性的2帧进行图像处理分析;根据碎片云图像特点以及碎片运动特性,提出了一种改进的碎片二次特征匹配算法,该方法包含碎片...     

Experimental and numerical study of the impact behavior of SMC plates

Steel components absorb impact energy by plastic deformation whilst composite materials absorbing it by damage mechanisms such as fiber debonding, fiber fracture, and matrix cracking. Therefore, in order to properly substitute metal components with composite ones in industrial applications, the impact property of composite materials must be well known. In this study, the impact behavior of sheet molding compounds (SMC), which is widely used in automobile industry due to its relatively low cost and high productivity, was examined both experimentally and numerically. In order to investigate the impact behavior of SMC, an experimental study was carried out by setting up a drop weight impact test system. Using this system, the dissipated impact energies of SMC flat plates were measured to investigate the influence of the mass and shape of impactor, initial velocity, and specimen thickness on the impact behavior.

For numerical predictions, a modified damage model for SMC was developed and adopted in the user defined material subroutine of the commercial simulation program LS-DYNA3D. For the sake of improving efficiency of impact simulations, the SMC material property was determined in consideration of the local differences of the fiber volume fractions. The dissipated impact energies under various conditions and the reliability of the developed impact simulation process were examined through comparisons of the predicted data with the experimental results.

From this comparison, it was found that, in the scope of current study, the specimen thickness is the most important parameter that should be considered in the design of SMC components for the aspect of impact behavior.