Search

Aims Out-of-hospital cardiac arrest (OHCA) affects around 1/1000 person-years. Following return of spontaneous circulation (ROSC), the patient can manifest neurological impairment. A targeted temperature management (TTM) protocol is recommended to prevent hypoxic–ischaemic brain damage in patients with coma after cardiac arrest. Neuro-prognostication remains substantial for the prediction of clinical outcomes. To study clinical characteristics, overall survival, and neurological outcome of patients with Glasgow Coma Scale (GCS) Methods Observational retrospective study evaluating all patients with OHCA of presumed cardiac cause and with GCS Results One-hundred and five patients out of 107 patients initially selected were included in the study (77% male, mean age 67 years). At 30 days, mortality was 41% and 53% of patients had a poor neurological outcome (Cerebral Performance Category, CPC, 3–5). Sixty-nine patients (66%) underwent TTM. In regard of the circumstances of OHCA, index event in a private place [OR = 3.12 (1.43–7.11), P = 0.005], ineffective rhythm changes during resuscitation manoeuvres [OR = 2.40 (1.05–5.47), P = 0.037] and a greater amount of adrenaline administered during resuscitation [OR = 1.62 (1.27–2.06), P Conclusions In our population, at 30 days after cardiac arrest, survival rate and the rate of good neurological outcome were comparable to those of the major international registries and studies. Even though patients treated with TTM did not demonstrate significant differences in terms of neurological outcome, this might be related to study-sample size and patient selection. Results in the literature are still controversial on this topic. The MIRACLE score showed a good performance, making it suitable for clinical use in our population. Similarly, the proposed multivariate models are potentially useful for the elaboration of simple and effective prognostic scores in neurological risk stratification.

CFX software is used to simulate different hybrid rocket configurations, applying liquid N2O as the oxidizer and paraffin as the fuel. This work is the prosecution of a previous paper analyzing liquid injection in a lab-scale hybrid rocket. It is focused on the formulation of the most suitable simulation technique to represent another type of liquid injector, compared with the one described in the previous paper. It also aims at extending the computational fluid dynamics simulation approach to hybrid rockets of larger scales. To validate computational fluid dynamics output, experimental results coming from both a laboratory scale and an increased-scale engine have been used. The different geometries studied include an increased-scale engine with a cylindrical grain having no diaphragm, the same rocket with a one-hole diaphragm inside the fuel grain, and a lab-scale rocket with a one-hole diaphragm. Simulations are steady state, and combustion derives from a single-phase chemical reaction. Liquid injection...

The study of hybrid rocket transient behavior is a very important issue in order to analyze instabilities and develop throttleable motors. In this paper, an unsteady numerical model is presented. The model is composed of three submodels linked together: 1) zero-dimensional model of the combustion chamber, 2) one-dimensional (1-D) radial model of the fuel grain, and 3) zero-dimensional model of the feeding system. The first model simulates combustion chamber dynamics. The one-dimensional radial model of the fuel grain includes both standard polymeric and liquefying propellants characterized by a melting layer. The fuel block models the heat exchange to the wall, the heat propagation through the solid/liquid phase, and the fuel surface evaporation/entrainment. Finally, the injection system block simulates the unsteady behavior of the feeding line and droplets break-up-evaporation dynamics. In this paper, all the blocks are presented together with their validation versus analytical test cases. For each block...

This paper presents a research work estimating the impact of propellant sloshing on the pointing-stability of the EUCLID satellite, for a pre-design phase of the project. The analysis is carried out by means of computational fluid dynamics. Two tank fill ratios have been considered: 22% and 50%, with an elastomeric and deformable bladder. To support the analysis, two different test cases have been created for the validation of the numerical approach. The first case is related to the analysis of a cylindrical tank with no bladder, and its importance consists in the determination of the correct computational fluid dynamics setup for sloshing simulations. For this test case, sloshing frequency is compared between computational results and the results of the analytical models available. The second test case is created to verify the procedure adopted for bladder simulation; therefore, it consists in the numerical simulation of a bladdered tank. For this case, the available experimental results have been used for a comparison with the output of the computational fluid dynamics analysis. This work is the first aiming at representing bladdered tanks using computational fluid dynamics without recurring to fluid–structure interaction techniques. Moreover, it proposes a new method for the calibration of the traditional sloshing lumped model, based on computational fluid dynamics, which could reduce development costs of experiments.

Computational fluid dynamics has been applied to the simulation of hybrid rockets using O2 as the oxidizer and hydroxyl-terminated polybutadiene or high-density polyethylene as the fuel. Simulations have been carried out using predefined fuel and oxidizer mass flow rates, and calculating the fuel mass flow rate as a function of the wall heat flux. In this second case, the calculated regression rate has been compared to the average value derived from the reference experiments; no tuning coefficients have been introduced. Computational fluid dynamics results are compared with the corresponding experiments in the literature. In those tests where regression rate is calculated as a function of the wall heat flux, the fuel grain is divided into segments and its average value is plotted as a function of the average oxidizer mass flux in the combustion chamber. Computational fluid dynamics regression rate is underestimated by 30% for high-density polyethylene and 50% for hydroxyl-terminated polybutadiene if only ...

A study on vortex injection in hybrid rocket engines with nitrous oxide and paraffin has been performed. The investigation followed two paths: first, the flowfield was simulated with a commercial computational fluid dynamics code; then, burn tests were performed on a laboratory-scale rocket. The computational fluid dynamics analysis had the dual purpose to help the design of the laboratory motor and to understand the physics underlying the vortex flow coupled with the combustion process compared with axial injection. Vortex injection produces a more diffuse flame in the combustion chamber and improves the mixing process of the reactants, both aspects concurring to increase the c* efficiency. A helical streamline develops downstream of the injection region, and the pitch is highly influenced by combustion, which straightens the flow due to the acceleration in the axial direction imposed by the temperature rise. Experimental tests with similar geometry have been performed. Measured performance shows an incr...

This paper reports a theoretical and numerical study aimed at increasing the operating efficiency of two-stage light-gas guns by appropriately changing their working conditions. In particular, a method is presented for increasing the projectile speed without any rise of the maximum breech pressure. The classic design theory of two-stage guns starts from the assumption that the highest velocity is reached in a gun which constantly maintains the maximum acceptable pressure at the base of the projectile during the full launching time. The main drawback of this working condition is that it may require an unfeasible rise of the gun maximum pressure, especially when very high muzzle speed is requested. To overcome this limitation, a new reference case different from the constant-base-pressure one is presented, based on a novel gas-dynamics solution that can be expressed in exact form if losses are not accounted for. According to such an approach, it is theoretically shown that the projectile base-pressure can be appropriately shaped (i) to improve the final speed without increasing the breech pressure or, in other terms, (ii) to achieve a given muzzle velocity with reduced maximum gas pressure. The analytical application of the new gas-dynamics condition showed the capability of obtaining a 1 km/s velocity improvement with no increase of the breech pressure or, alternatively, a pressure reduction up to 30% with no penalty on the model final speed. A numerical verification of the calculations was performed through the CISAS light-gas gun full numerical model, which includes real effects such as friction losses and heat transfer. Finally, an experimental verification of the numerical test case was attempted and a speed augmentation of 0.8 km/s with no increase in the breech pressure was confirmed in laboratory, highlighting the agreement with numerical predictions.

Impact processes influence the surface evolution of the solid bodies in the Solar System, as asteroids or comets. The better understanding of these phenomena can improve the interpretation of remote sensing data from forthcoming missions. In order to complement and to extend the available data on hypervelocity impacts on porous targets to ranges of velocity and physical conditions not yet explored, we focused our work on catastrophic fragmentation and cratering processes onto porous targets by means of high velocity impact experiments using a two-stage light-gas gun located at the impact facility of CISAS “G. Colombo” of the University of Padova. Tests have been performed on targets of different materials, e.g. glass ceramic foam, natural pumices, water ice, and different porosity (with density in the range from 0.35 to 1.07 g/cm 3 and porosity from 2% up to 65%). Results have been analysed and compared to published data. In particular, cratering morphology is studied as function of projectile velocity and density and then it is compared to empirical trend reported in Kadono [Kadono, T. Hypervelocity impact into low density material and cometary outburst. Planet. Space Sci., 47, 305–318, 1999.] with good agreement. Our tests results revealed that porosity seems to induce a rapid attenuation of the shock affecting energy propagation inside the target; craters generated by hypervelocity impacts on porous targets are smaller and much deeper than those of non porous targets. Fragments distribution resulting from disruption tests is analysed and the largest fragment mass has been studied as function of the collisional specific energy; results is compared to those of paper by Ryan et al. [Ryan, E.V., Davis, D.R., Giblin, I. A laboratory impact study of simulated edgeworth-Kuiper belt objects. Icarus 142, 56–62, 1999.]. Furthermore, numerical simulations have been performed by using Smooth Particle Hydrodynamics (SPH) and Lagrangian grid technique in order to test some available numerical models for a forthcoming more accurate numerical analysis that will be validated with the comparison with experimental results.

In this paper, CFX is used to simulate different hybrid rocket configurations applying liquid N2O as the oxidizer and paraffin wax as the fuel. This work is intended as the prosecution of a previous study about hybrid rockets with diaphragms of different geometries inside the combustion chamber, where N2O was injected in a gaseous phase. In this work, liquid injection is introduced, together with droplets vaporization and their coupling with the Eulerian gas phase, in terms of both heat and momentum exchange. The main objective is the description of the numerical models to be applied when liquid is injected. To validate computational fluid dynamics output, experimental results coming from a laboratory scale engine have been used. The different geometries studied include an engine with a cylindrical grain having no diaphragm and the same rocket with a four-hole diaphragm at 24% of the grain length. The simulations are steady state, and combustion derives from a single-phase chemical reaction. Liquid inject...

This paper describes computational fluid dynamics applied to the analysis of a hybrid rocket motor with a diaphragm in the combustion chamber to enhance rocket performance. This work follows the last author’s experiments: An engine was tested with nitrous oxide and paraffin wax as propellants. Several of the tests have been used as references for numerical simulations. The following approximations have been made: steady-state conditions, eddy dissipation model with one-step reaction, gaseous injection of fuel and oxidizer, and no droplets entrainment (typical of a paraffin grain). First of all, a single geometry without a diaphragm has been analyzed with different turbulence models (k-ω, k-ω shear stress transport, k-e, k-e renormalization group). It has been shown that the k-ω model predicts a lower flame temperature and chamber pressure than the k-e model. Then, five geometries have been studied to compare two different types of diaphragms (one hole and four holes) in two positions (24 and 33% of the to...

This paper presents a research work estimating the impact of propellant sloshing on the pointing-stability of the EUCLID satellite, for a pre-design phase of the project. The analysis is carried out by means of a simplified lumped parameter model describing the fluid system. This model allows the identification of forces and angular moments due to liquid movement under different acceleration conditions, and for different fill levels of the reservoir. The lumped model is calibrated by means of CFD simulations to determine the correct values to be attributed to the different coefficients. For this work, two tank fill ratios have been considered: 22 and 50%, with an elastomeric and deformable bladder. The results of the lumped models are compared with the CFD output obtained using the same acceleration profile, and corresponding to the same fill level. The study shows that using CFD to calibrate the lumped parameter model is a promising technique, which can be applied at a preliminary design stage, to reduce development costs and time. This work is intended as the prosecution of the first study about sloshing by the same authors, which was aimed at representing bladdered tanks using computational fluid dynamics without recurring to fluid–structure interaction techniques. Here, a new method for the calibration of the traditional sloshing lumped model, based on computational fluid dynamics instead of real experiments, is proposed and applied to a specific fluid system. This procedure can result in a consistent reduction of the experiments costs.

The EU FP7 SPARTAN project (SPAce exploration Research for Throttleable Advanced eNgine) aims at developing a highly throttleable propulsion system (10:1 thrust ratio), to perform soft and precise landing in planetary environments; the innovative technology relies on a hybrid engine burning HTPB and decomposed 87.5% H2O2 with vortex injection. In this framework the University of Padua (UPD) is responsible for developing a prompt and reliable CFD tool to support both design phases and internal ballistic analysis of the hybrid engine. To validate numerical results UPD has developed its own engine and ground testing are still ongoing to characterize hybrid performance in deep throttling. The issues related to handling and storing high concentration hydrogen peroxide have been overcome integrating the hybrid motor with a dedicated gas generator to reproduce 87.5% H2O2 in dissociated condition. The mixture produced inside the gas generator is then injected as oxidizer into the combustion chamber of the hybrid motor. This paper presents the main results of the optimization tests campaign performed on the integrated gas generator, to assess the injected hot mixture characteristics and ignition capability of the HTPB fuel.

The SPARTAN (SPAce exploration Research for Throttleable Advanced eNgine) project is targeted at the development of a soft-lander demonstrator for Mars landing, based on throttleable hybrid rocket engine propulsion technology. Within the project, CISAS “G. Colombo” is in charge of the development of a new advanced CFD code for the hybrid rocket motor unsteady simulation, of other CFD simulations conducted with commercial software to investigate the fluid dynamics of the rockets developed by NAMMO, and of ground fire tests to be conducted at lab-scale. Thus, for the CFD simulations, an Open Source code (OpenFOAM) and a commercial code (Ansys CFX) have been used. The flow field has been investigated in its different aspects, to understand the physics behind the processes of injection and combustion. The main approximations applied to this study when analyzing the flow with commercial codes are steady state analysis and fixed chemistry in the combustion chamber. Both vortex and axial injection have been simulated by means of the CFD: an axial injector has been used for the CISAS lab-scale rocket motor and a vortex injector for the test cases reproducing NAMMO’s configurations. Regarding vortex injection, the results highlight that it improves the turbulent mixing between the reactants and pushes the flame near the wall, enhancing the heat flux transmitted and thus the regression rate, thanks to its strong helical flow and centrifugal effects. Further efforts have been oriented to take into consideration regression rate dependency on the wall heat flux: besides injecting a fixed and pre-established fuel mass flow, the latter has been coupled to heat exchange. This study is in a preliminary phase, but the first results show that CFD regression rate as a function of the oxidizer mass flux follows the same trend as predicted by: , with a and n experimentally estimated by NAMMO. Concerning OpenFOAM numerical solver, it provides a transient solution of the flow field, but uses corresponding boundary conditions compared to commercial software, in order to allow a full comparison of the steady state results obtained. The study performed with OpenFOAM is a work in progress. The starting point has been a comparison with the results obtained using commercial CFD, to validate the open source solver and boundary conditions. This comparison has showed that OpenFOAM can predict the experimental results as well as commercial software does. At the moment, a feasibility study is being conducted in order to develop a custom and flexible boundary condition, able to calculate regression rate as a function of the wall heat flux independently on the specific oxidizer/fuel combination. Other studies are also necessary to limit the computational time required by this open source tool.

In recent years paraffin wax has been introduced as a fuel for hybrid rocket motors. Its fundamental characteristic is a higher regression rate compared to classical polymeric fuels due to entrainment of liquid droplets from the fuel surface. For this reason paraffin wax is the main fuel used by CISAS in its research experiments. In order to study paraffin-based fuels behavior a 1D transient code of the fuel grain has been implemented. The first part of the paper describes the equations used by the numerical model and its validation. The second part of the papers deals with the phenomenology of supercritical entrainment. It is shown that above the critical pressure fluids behave in a different way respect to liquids. Three hypotheses have been done in order to describe supercritical entrainment. Finally the regression rate predictions of the model are compared with the experimental results. It is shown that the slope of the regression rate curve is strongly related to the entrainment law exponents because the vaporization regression rate is small compared to the entrainment part. The closure problem of droplet entrainment is described, highlighting the need for a physical based link between surface temperature, droplets temperature and vaporization temperature.

This paper deals with an experimental and numerical project intended to study fully tangential vortex injection in a hybrid motor of 1kN class. Due to the knowledge of the CISAS Hybrid Team, the choice for oxidizer has been pressurized nitrous oxide, while paraffin wax has been used as fuel. This investigation follows a previous project where a mixed axial/vortex device has been tested: also if an increase in performance has been observed, a non-uniform grain consumption showed an issue of such device. The work starts with a mission scenario that gives the design drivers for the subsequent preliminary design, performed with an iterative, transient and 0D numerical code. A successive step is the detailed design of the combustion chamber and test bed. The aim of the study was to investigate three objectives: vortex injection and a comparison with axial; throttling behavior at fixed mass flows (reduction of 75% and 50% of oxidizer flow); combustion chamber performance changing its configurations, in particular using inside a mixer. Performance parameters taken into account were chamber pressure oscillations, combustion efficiency and regression rate. In the preliminary phase two different issues have been discovered and solved: the first regards a chamber pressure behavior variation, linked to a too long postchamber; the second is referred to pressure pikes in the ignition phase, solved using a longer prechamber and a different ignition configuration. It has been shown that vortex injection lowers the chamber pressure oscillations respect to axial case from more than 7% down to 4%. Moreover, regression rate has been increased of 41%, and the a coefficient of its law up to 67% from axial. This last value, indeed, shows a constant behavior throttling down the oxidizer mass flow. The increase is due to the higher wall heat flux in the grain surface, given by the higher velocity and thermal gradient of the fluid. Combustion efficiency has been increased with vortex injection given by the higher turbulence flow that enhances the mixing of the reactants. Axial case showed a value of 76% of this last parameter, while vortex case went up to 90%. Coupling of injection and mixer (a diaphragm-like device inside combustion chamber) increases this value up to 96%. A cost analysis has been performed for this project, showing that hybrid propulsion is a low cost technology.

One of the most effective measures for a rapid return of military submarines to the surface in case of life threatening, hazardous or potentially catastrophic failures is the emergency blowing of the forward main ballast tanks. This can be achieved with various systems. Gas generator based modular emergency blow systems show several advantages compared to designs, which use highly compressed air in terms of specific gas work, mass flow and gas storage density. Operational gas generators for emergency blowing use either a liquid hydrazine monopropellant or a solid propellant consisting of GAP and strontium nitrate for gas generation. An alternative hybrid propellant gas generator is proposed to avoid safety issues, which both operational gas generators types possess. Also the relative free selection of the blow gas composition motivates the use of a hybrid propellant combination. A mass model for a hybrid gas generator is developed using design elements of the hydrazine gas generator and applied to a number of different hybrid propellant combinations. The hybrid propellant mass is determined by use of the assumption that the volume reduction due to heat losses, water condensation and gas dissolution in the ballast tank can be compared for a blow gas that is generated from a solid or a hybrid propellant. The effectiveness of the modeled hybrid gas generator is assessed amongst others by blow system mass, system volume and required gas generator quantity. Comparing these preliminary results to the operational solid and liquid propellant gas generator shows that even a nonoptimized hybrid propellant gas generator system is very probably competitive in terms of mass and volume.

Combustion efficiency in standard hybrid rocket motors is usually low. This is due to the poor mixing of the core oxidizer flow with the gasi fied fuel from the walls. Motor scale increase usually worsens this issue. Large and volu metric inefficient post combustion chambers are common devices used to improve efficiency, while mixers are somewhat more complex, but more effective devices. The addition o f a mixer at grain aft-end was originally suggested by Marxman [1], but literature data on mixers, and in general on techniques to improve combustion efficiency is lacking. The resea rch activity herein presented aim at further investigate this issue: a mixer placed in f uel grain was selected as efficiency enhancing device. This work starts from experiments performed by Dr. Grosse [1] and more testing, both at lab scale and at increased scale h ave been carried out. Simple 1-hole diaphragms fixed at 25% of the grain length were used. Centre-hole diameter has been optimized. Moreover, CFD simulations have been carried out to further understand the enhanced mixing phenomena. N2O was selected as oxidizer and as fuel the same par affin Sasol Wax 0907 mixture used by Dr. Grosse was chosen. Lab scale experiments showed that combustion efficiency is raised from 76% (without diaphragm) up to 95% using diaphragms. Also good agreement has been found with CFD simulations. At increased scale (three times the thrust of lab scale), the referenc e combustion efficiency was 80%, and it was raised at 94% with the addition of a diaphragm. Als o combustion stability was enhanced. Another plus of the diaphragm on post-chambers and aft-end mixers is to enhance regression rate downstream it. At lab scale and wit h the smaller diaphragm regression rate was increased up to +90% (4.5 mm/s) compared to literature data without diaphragm. At increased scale the increase was of +65 % (4.2 mm/s). Further research is needed, but these experiments showed that diaphragms can be used to design compact and efficient single-port and paraffin based hybrid rocket motors.

A study on vortex injection in hybrid rocket motors with nitrous oxide as the oxidizer and paraffin as the fuel has been performed. The investigation followed two paths: first of all, the flow field was simulated with a CFD code, and then burn tests were performed on a lab-scale rocket. The CFD analysis had the dual purpose to help the design of the lab motor and to understand the physics underlying the vortex flow coupled with the combustion process. Numerical analysis was focused on the comparison with axial injection. Vortex injection produces a more diffuse flame in the combustion chamber and improves the mixing process of the reactants, both aspects concurring to increase the efficiency of the motor. A helical streamline develops downstream the injection region, and the pitch is highly influenced by combustion, that tends to straighten the flow due to the acceleration imposed by the temperature rise to the axial velocity component. The tangential velocity, on the contrary, is far less influenced by this effect. Experimental tests with the same chamber geometry have been performed with both pressurized and self-pressurized oxidizer. Measured performances showed an increase in regression rate up to 51% and a combustion efficiency that rises from values lower than 80% in the axial injection configuration up to more than 90% with vortex. Moreover, a reduction of the instabilities in the chamber pressure has been measured. Issues requiring further investigation concern the motor exhausts: both experimental and numerical analyses showed that there is a residual tangential velocity component in the plume; this, coupled with a noise suppressor system downstream the nozzle in the test apparatus, showed severe instabilities in the vortex configuration thrust measurements, not reported in chamber pressure burn data and not affecting axial injection.

In order to analyze the propagation of shocks due to hypervelocity impact of micrometeoroids and space debris on spacecrafts, it was necessary to analyze the high-frequency shock-propagation-dynamic on complex structures representative of real spacecraft structure. Objective of this research is the GOCE satellite due to its highly accurate accelerometer very sensitive to the micro-vibration environment. After a detailed analysis of the most-probable shock-propagation-path within the satellite, several representative targets have been designed. Then an extensive test campaign has been conducted on these targets exploring a wide range of impact conditions. As a result, a database was established which correlates the impact conditions in the experimental range (0.6–2.3 mm projectiles at 2.5–5 km/s) with the shock spectra on selected locations on various types of structural models. The behaviour of structural joints was also analyzed under shock conditions that are not normally reproduced by pyroshock testing. The database represent a fundamental tool in order to validate the numerical analysis that will be used to assess the vibration environment, due to in-orbit micrometeoroids and space debris at the accelerometer location. This paper presents the test procedure applied and the general results achieved.

To design a launcher system competitive to current launcher technology high combustion chamber pressure is required. In this paper two types of turbo-pumps technologies have been considered in relation to hybrid propulsion systems: (i) spill of hot gas from the combustion chamber, (ii) regenerative cycle. To evaluate effectiveness of each technology in relation with the selected mission profile a global hybrid thruster model has been applied. The model performs both internal ballistic calculation and provides a preliminary size of the thruster structure in term of masses and dimension. Effectiveness of pressurization technology has been evaluated for the following mission profiles: Ariane booster, first and second stage, VEGA first and second stage, Soyuz first second and third stage. For each configuration a comparison among hybrid and state of the art propulsion technology has also been done based on propulsion system mass, size, safety issues and reliability.

Among all possible dangers, hypervelocity impacts on structures produce also a disturbance field, which results from the superimposition of vibrations originating on the impact point and then reflected at the target boundaries. Such disturbance environment is composed by waves of different amplitudes, frequency contents, speeds and directions of propagation. The aim of this paper is to characterize this complex environment, through the identification of its fundamental constituents using the Wavelet Transform analysis method, useful for studying transient phenomena, including wave propagation. Common signal analysis tools, like Shock Response Spectrum (SRS), cannot provide a good description of the physical behavior of such waves, nor they can differentiate between them since their frequency decomposition does not retain any time information. On the contrary, Wavelets associate a time information to the frequency content: each wave can be characterized by its frequency band and arrival time. Starting from both numerical and experimental acceleration data from impacts on aluminum plates and aluminum honeycomb sandwich panels, Wavelet analysis was employed to identify symmetric and antisymmetric waves. Moreover, reflections and dispersion phenomena were observed, leading to wave distortion due to different speeds of propagation of wave trains characterized by different frequency contents.

This paper presents the results of a set of experiments aimed at discovering the main features of impact-induced vibrations on all-aluminum honeycomb sandwich panels, representative of the GOCE satellite's top floor, which is exposed to the orbital debris environment. The activity focused on the characterization of the vibrations induced in the vicinity of internal payloads by hypervelocity impacts occurring on the vehicle's external shell. More than 30 tests were realized by launching 0.8–2.3 mm aluminum projectiles in the velocity range 4–5.5 km/s on targets with tri-axial accelerometer assemblies mounted on both the front and rear face of the panel, at a nominal distance of 150 mm from the impact point. It was found that a hypervelocity impact produces in both the front and rear side of the sandwich panel a vibration environment which can be described through the shock response spectrum (SRS) of three different types of waves that can be distinguished on the basis of the acceleration direction: out-of-plane, in-plane longitudinal and in-plane shear. The influence of projectile mass and velocity on SRS appeared to vary with frequency, with the most significant difference in the range between ∼103 and ∼104 Hz. The results of whole experimental set were used to derive an interpolation law through standard techniques of nonlinear fit. The empirical equation obtained makes it possible to predict the near-field vibration environment produced by hypervelocity impacts with debris having given size and velocity, reproducing all the test data with an average uncertainty of ±6 dB.

This paper presents an overview of the hypervelocity impact test campaign ongoing in the frame of the ESA contract “spacecraft disturbances from hypervelocity impact”. The project aims at analyzing the propagation of shocks due to hypervelocity impacts from the external shell of a spacecraft to its internal components. The object of the study is the GOCE satellite, which has been recognized to be very sensitive to small disturbances because of its payload that has been designed to measure even very low acceleration levels. In the first step presented hereafter, the test campaign has been focused on the qualification of the background environment inside the impact chamber and on the determination of the vibration levels induced by perforating and non-perforating hypervelocity projectiles on simple aluminum plates. The results currently obtained and a preliminary data analysis will be presented in the following.

This paper describes the SPONGE experiment (Sounding rocket Propellant OrieNtation microGravity Experiment) developed at CISAS in collaboration with Thales Alenia Space Italy. The aim of this experiment is to provide the data for the validation of the CFD code created to study and design propellant management devices. SpongeCompressibleFoam is a code based on the OpenFOAM Platform, written at CISAS to simulate propellant management devices. The latter are passive static metal structures used in rocket tanks to control propellant behavior; they work using surface tension to ensure gas-free liquid delivery to the tank outlet. SPONGE flew on board the REXUS 9 ESA/SSC/DLR sounding rocket in February 2011, it is based on two counter rotating plates: (i) the experimental plate, on which the control equipment and a polycarbonate tank containing the sponge test-sample are placed and (ii) the balancing plate, rotating in the opposite direction with respect to the experimental plate and ensuring no momentum transfer to the rocket. The system rotates at four different angular velocities, allowing the study of the sponge and its retention capability under different centrifugal forces. The design and results are presented.