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Application of the method of removal of space debris objects by an ion beam - the IBS method - seems to be the most effective for cleaning the GEO region. This is due to the low energy consumption for transfer one object in a safe orbit delta-V≃11 m/s and the possibility of removal several objects by one spacecraft. The service spacecraft for this purpose can be developed on the basis of telsat technology, taking into account its functions. Such a spacecraft should contain an ion beam injector to act on the removed object and an electric propulsion system, which serves to compensate for the ion injector thrust and to control the spacecraft motion both during transportation of the removed object and during interorbital flights. The development of a weakly diverging ion beam injector, determination of the composition and structure of an electric propulsion system, ballistics of the optimal sequence of interorbital flights, and development of an algorithm for controlling a virtual cluster: spacecraft-removed object are discussed in the article. In the majority of works on IBS, commercial ion thrusters are considered as an ion injector and engines. However, ion beams from ion thrusters have a half angle of divergence of 120-150. When propagating in outer space, the ion beam expands irresistibly due to electron thermal effects. As a result, the distance at which the spacecraft must follow the removed object during its transportation for effective impulse transmission to it should be small (estimated, about 7 m). Taking into account the dimensions of both objects and the uncontrolled movement of the removed object relative to its center of mass, the object transportation operation carries noticeable risks. The results of an experimental study of a prototype of an ion injector with an initial half angle of divergence of 20-40 are presented. The use of such an injector would provide a distance between objects of 15–20 m and reduce the risk of their collision. In connection with IBS, various aspects of spacecraft control during transportation of an object with one, two or more engines as part of the electric propulsion system are considered. In the conceptual design of the spacecraft presented in this work, the electric propulsion system contains two stationary plasma thrusters, each of which is mounted with the possibility of independent rotation respect to two mutually perpendicular directions. Spacecraft motion control, including thrust vector control, is carried out by means of the thrusters coordinated rotation. The characteristics of the spacecraft and its subsystems intended for the removal of several objects are conceptually considered. For example, parameters of the mission for the removal of up to 10 objects from the NORAD catalog from the GEO region are given. In the future, using the result obtained, the control of SC of this scheme when IBS implementing in the low-Earth orbit will be considered.

A mathematical model is developed to characterize the progressive time-evolution of a fragmenting incandescent object. The objective of these models is to provide a spatio-thermal footprint of the fragmentation field, which can be useful to guide fire safety rules in manufacturing workplaces, as well as to estimate fire hazards. Ascertaining the time-evolution of the temperature of the fragments is quite difficult to measure experimentally, which motivates the model development. Initially, analytical models based solely on ballistics, which provide qualitative trends, are developed to provide insight into the fundamental ratios that govern safe operating conditions. Thereafter, rapid numerical spatio-thermal models, which provide quantitative information, are then developed, based on particle methods. The model uses the released energy from the initial blast pulse to provide the starting kinetic energy of the system of particles and then numerically computes the trajectory and thermal state of the fragments under the influence of Numerical examples and provided and extensions to high-fidelity are discussed.

The French-German Research Institute of Saint Louis proposed an innovative concept of a microair vehicle (MAV). The gun-launched MAV (GLMAV) is a new and original type of discreet and robust miniature hybrid system of about 1 kg for a very quick observation mission by air. The GLMAV in a projectile configuration is launched by a portable dedicated tube weighing less than 10 kg. After a ballistic flight, the GLMAV is transformed into a miniature drone with automatic deployment of coaxial counter-rotating rotors. The GLMAV in drone configuration is slowed down to be able to execute hovering or/and maneuvering flights. In few seconds, aerial observation down and forward can begin in order to capture information, because it is equipped with an embarked vision system with real-time image transmission. This paper is first focused on the conception of the launcher and of the platform in order to incorporate all the features coming from multiple disciplines. This hybrid system conception is linked to an iterative process between all fields of research such as aerodynamics, interior and exterior ballistics, mechanics, electronics, optronics and signal transmission, guidance, navigation, and control. All theoretical studies presented in this paper have then converged toward the manufacturing of the dedicated ballistic launcher and of the GLMAV platform. The latter is now ready for the autopilot implementation in the near future.

A laboratory gun that uses compressed gas as its propellant was designed and tested for the study of small projectiles traveling in soils with initial velocities ranging from 50 to 200 m/s. The gun employed the use of air or helium with an electrically triggered, pneumatically piloted, solenoid valve. The triggering system was designed to provide remote activation. A photo gate speedometer was utilized to measure the exit velocity of each projectile. Details of the electro-pneumatic control system are presented in this paper along with the design of the gun assembly and its subsystems. The effects of different design parameters, including muzzle length, projectile mass, propellant type, and volume of compressed gas utilized, on projectile velocity were investigated. Statistical analysis of the gun performance is presented. The gun is currently being used to visualize the fundamental physics of rapid earth penetration in soils, using transparent soils.

The aim of this paper is to determine the velocity, energy, maximum range and distance at which pellets fired from an air rifle of kinetic energy below 17 J can pose a threat to unprotected human skin. Doppler radar equipment and exterior ballistics software were used in this examination.

We present an attitude estimation method for self-guided ballistic munitions from onboard magnetometers. An adaptive feedback filter is used to remove actuator-induced magnetic disturbances whose spectrum overlaps the signal of interest. A state-augmented extended Kalman filter (EKF) is used to estimate roll rate and sensor bias from the filtered data. Experimental results demonstrate that this method is able to accurately estimate the roll without phase delay inherent in other methods.

In this paper, a new proportional-navigation guidance law, called retro-proportional-navigation, is proposed. The guidance law is designed to intercept targets that are of higher speeds than the interceptor. This is a typical scenario in a ballistic target interception. The capture region analysis for both proportional-navigation and retro-proportional-navigation guidance laws are presented. The study shows that, at the cost of a higher intercept time, the retro-proportional-navigation guidance law demands lower terminal lateral acceleration than proportional navigation and can intercept high-velocity targets from many initial conditions that the classical proportional navigation cannot. Also, the capture region with the retro-proportional-navigation guidance law is shown to be larger compared with the classical proportional-navigation guidance law.

Radar is assumed to be tracking a cannon launched projectile travelling in a 2 dimensional drag environment. Extended Kalman Filter (EKF) is explored for the estimation of position and velocity of the projectile. Simulation is carried out using MATLAB and the results are presented. Finally, based on the results, EKF is recommended for tracking the projectile.

With more than 1 trillion artificial objects smaller than 1 μm in low and geostationary Earth orbit, space assets are subject to the constant threat of space debris impact. These collisions occur at hypervelocity or speeds greater than 3 km/s. In order to characterize material behavior under this extreme event as well as study next-generation materials for space exploration, this paper presents a unique two-stage light-gas gun capable of replicating hypervelocity impacts. While a limited number of these types of facilities exist, they typically are extremely large and can be costly and dangerous to operate. The design presented in this paper is novel in two distinct ways. First, it does not use a form of combustion in the first stage. The projectile is accelerated from a pressure differential using air and inert gases (or purely inert gases), firing a projectile in a nominal range of 1–4 km/s. Second, the design is modular in that the first stage sits on a track sled and can be pulled back and used in itself to study lower speed impacts without any further modifications, with the first stage piston as the impactor. The modularity of the instrument allows the ability to investigate three orders of magnitude of impact velocities or between 101 and 103 m/s in a single, relatively small, cost effective instrument.

In the first part of the paper [Naik D, Sankaran S, Mobasher B, Rajan SD, Pereira M. Development of reliable modeling methodologies for fan blade-out containment analysis. Part I: experimental studies. Int J Impact Eng, in press], details of the experiments to characterize the behavior of dry fabrics including Kevlar®49, and ballistic tests involving the fabric were presented. In this second part of the paper, we discuss the development and verification of a constitutive model for dry fabrics for use in an explicit finite element program. The developed constitutive model is implemented as a user-defined subroutine in LS-DYNA, a commercial finite element program. It is then used to simulate a suite of ballistic tests [Naik D, Sankaran S, Mobasher B, Rajan SD, Pereira M. Development of reliable modeling methodologies for fan blade-out containment analysis. Part I: experimental studies. Int J Impact Eng, in press] that replicate conditions seen in an engine fan blade out (FBO) event. A qualitative and quantitative comparison of the deformation shape of the fabric containment system and a quantitative comparison of the absorbed energy are carried out. Results indicate that the developed constitutive model provides a very encouraging start in modeling these high-speed events and could provide as Federal Aviation Administration (FAA) desires, an attractive complement to full-scale engine FBO tests.

Ballistic imaging is the name applied to a category of optical techniques that were originally developed for medical applications. Recently, ballistic imaging was adapted to acquire instantaneous images of the liquid core inside atomizing sprays; a region that has been heretofore inaccessible to spray researchers. An important difference between spray research and the medical imaging problem is the need for high fidelity single-shot (within 10 ls) imaging in a spray whereas stationary tissue images can be averaged. Transient ballistic imaging diagnostics have been used to reveal details of the primary breakup process in a LOX injector, a turbulent water jet, a water jet in cross-flow, a transient diesel fuel spray, a rocket fuel injector, and an aerated spray. This paper briefly discusses various methods for imaging the liquid core, it introduces ballistic imaging and provides specific examples, it describes detailed studies of photon transmission through dense media, and it then discusses incorporation of those results into a model for a ballistic imaging instrument that can evaluate and optimize various concepts.

The present paper constructs a rigid–plastic model to assess the effect of a soft nose on the perforation of metallic plates struck by a stubby projectile. A low-density nose-cabin-column, i.e., an assembly of seeker and guidance/control unit, is installed in the front of a real missile. It usually acts as an energy absorber and buffer in the perforation of metallic plates and notably affects the terminal results of projectile. The problem is equivalent to a blunt rigid projectile with a soft nose perforating a circular metallic plate. Effects of transverse shear, bending and membrane deformations on the perforation process are included in a rigid–plastic analysis. Especially the effect of an ahead structural response, induced by the soft nose impacting the target plate, is considered in the model. Analytical formulae for the ballistic performance are obtained for a range of plate thicknesses, which agree well with available experimental results.

Helmet designs have evolved over the last three millennia and the use of helmet may be as old as warfare itself. In 600 B.C. the Greeks in Sparta crafted single-piece helmet from bronze, which provided complete head protection, leaving only narrow slits in front for vision and for ventilation. Later in 250 B.C. the Romans developed several helmet designs, which included the round legionary's helmet and the gladiator's helmet, with broad brim and pierced visor, providing exceptional head, face and neck protection. The use of helmets in battle fields continued until the end of 13th century, which signaled a radical change in the emphasis on head protection. With the invention of gunpowder and the growing effectiveness of firearms over swords and spears, helmets gradually vanished from the battle scenes. The metal helmet that once protected against sword and arrows offers little protection against musket rounds. The United States Civil War provides an excellent illustration—in war soldiers donned cloth hats and caps, with little or no emphasis on head protection. In World War I, the helmet was reintroduced because it protected the head against metal-fragments of exploding artillery shells and indirect fire. The French, owing to General Adrian, were the first to adopt the helmet as standard equipment in early 1915. The British, the Germans, and then the rest of Europe soon followed. In World War I, the German helmet provided the best protection for soldiers. It was manufactured in at least two sizes, and different head contours were accommodated by an adjustable leather lining, which provided comfort and allowed for ventilation. The typical helmet used during that period was a hardened steel shell with an inner liner and weighed about 0.5–1.8 kg. Since then, helmets have been issued to troops in all military conflicts, marking the beginning of the development of modern military helmet. This paper presents the results from experiments and AUTODYN-3D® simulations on the ballistic impact of a KEVLAR® helmet. In the experiment, spherical projectile (~11.9 g), launched from a light gas gun, strikes the helmet with an impact velocity of 205 m/s. The interaction of the projectile with the KEVLAR® helmet is captured using high-speed photography. This helmet-projectile interaction is compared with that obtained from the AUTODYN-3D® simulation. Post-test damage photos from the experiments are also compared with those from the simulations. The response of the helmet from the simulations is consistent with those from the experiments. Also included in this paper are AUTODYN-3D® simulations on two ballistics test standards for KEVLAR® helmets. They are namely the NIJ-STD-0106.01 Type II and the V50 requirement of the US military specification for Personal Armor System Ground Troops (PASGT) Helmet, MIL-H-44099A. For the simulation on MIL-H-44099A, a fragment-simulating projectile (FSP) strikes the helmet with an impact velocity of 610 m/s. The simulation revealed that an impact velocity above 610 m/s is required to perforate the KEVLAR® helmet. For the simulation on NIJ-STD-0106.01 Type II helmet, the projectile is a 9 mm full-jacketed bullet with a striking velocity of 358 m/s. Results from the simulation show that the KEVLAR® helmet is able to defeat a 9 mm full-jacketed bullet traveling at 358 m/s.

Internal ballistics simulations of solid rocket motors have been conducted with the propellant grain's 3-D burning surface geometry described by a new minimum distance function approach and the internal flowfield represented by 1-D, time-dependent, single-phase compressible flow equations. The combustion model includes erosive burning and unsteady, dynamic burning corresponding to transient energy storage in the heated surface layer of the propellant. The integrated internal ballistics code (Rocballist) is used to investigate the role of these two burning rate augmenting mechanisms in solid rocket motor performance. Two tactical motors are used as test cases. Results indicate that dynamic burning can be the dominant factor in producing a short-duration ignition pressure spike in low-L* motors, particularly if the L/D ratio is not too large and the port cross section is nonrestrictive (e.g., center perforated grain). However, when L/D is large and the port cross section is noncircular in the aft section (aft fins/slots), erosive burning can take over in dominating the burning rate to the extent that an otherwise progressive pressure-time trace becomes regressive/neutral. That is, erosive burning can effectively prolong the initial pressure spike in some star-aft motors. The results also show that with sufficiently accurate models of dynamic burning and erosive burning, it is reasonable to expect reliable internal ballistics predictions with suitable simplified flowfield models, thereby realizing significant reductions in computation time compared with 3-D, multiphase reacting flow simulations.

Woven fabrics are widely used in soft body armour. The ballistic performance of a range of narrow fabrics was determined and compared to that of wider fabric panels. In contrast to conventional ballistic fabrics, it was found that the performance of narrow fabrics was highly sensitive to changes in fabric specification. Furthermore, it was shown that square weave fabrics may not necessarily be superior to unbalanced fabrics in narrow fabric form. It was also confirmed that narrow fabrics absorb more energy than wider fabric panels when undergoing direct ballistic impact, particularly when gripped in a two-sided configuration. Ways were sought of utilising these improvements in performance when the narrow fabrics are brought together to make full-width armour panels. Potentially, the observed improvements in performance could have been offset by the lines of weakness between adjacent fabric strips or by the weight of the structure required to grip on two sides. However, after investigating a number of potential solutions, significant improvements in the performance-to-weight ratios over simple fabric armour were achieved using an assembly of narrow fabrics attached to a lightweight composite frame.

A research program was funded by the European space agency (ESA) to improve and optimize the shields used to protect the manned elements of the international space station (ISS) against impacts of micro-meteoroids and orbital debris. After a review of existing shielding systems and after a series of light gas gun (LGG) experiments to screen interesting new materials and configurations, the research focused on shields with a metallic outer bumper, an intermediate stuffing and an inner metallic wall (representing the pressure shell of a manned spacecraft). Additional LGG experiments were performed on several configurations, with bumpers containing aluminum foam or made from titanium and aluminum super-alloys and with several combinations of stuffing materials. The comparison of the test results showed that ceramic cloth (Nextel) plus aramid fabric (both 2D and 2.5D Kevlar weaving) used as intermediate bumper gave a good protection compared to the overall area density requested. Configurations with by-layered aluminum foam bumpers (sandwich panels with asymmetric Al face sheets and a core made from Al foam) and Kevlar stuffing showed excellent resistance to normal impacts at about 6.5 km/s. However, the influence of material properties varying from batch to batch and threshold phenomena made ranking among the tested options rather difficult. The test campaign showed that it was rather difficult to improve over the already good ballistic performances of the debris shields developed by Alenia Spazio for the ISS manned elements. The by-layered Al-foam bumper and Kevlar stuffing configuration was selected for additional tests, including low velocity and oblique impacts, to develop ballistic limit curves.

There has long been a need to understand the impact response of explosive materials, and continual improvements result from the design of careful, well-instrumented experiments. This article summarizes details of the design and construction of a laboratory facility capable of launching projectiles at explosive targets at velocities up to 1500 m s−1. There are two types of experiment that are required. In the first, a gun launches a plate of great planarity at an equally flat target. This geometry is known as plate impact and a target loaded in this manner experiences a state of one-dimensional strain. This loading is accomplished by launching plane impactors onto targets aligned to micron tolerances, normal to the impact axis to less than 0.5 mrad of tilt. In the second, it is required to attain the ability to recover impacted explosive targets that have been loaded in one-dimensional strain for subsequent microstructural assessment. The development of this capability will be described in a subsequent pub...

The physical and chemical phenomena involved in the ballistic process, both in conventional guns and in electrothermal-chemical (ETC) guns, are very complex, so that optimization of gun performance requires a detailed understanding of the influence of different parameters. The initial stage of propellant ignition is critical to the rest of the process, in terms of both performance (muzzle velocity, repeatability) and safety. A better understanding of the ignition stage is thus very important in achieving improved performance. In this paper, we propose a new method and ignition sensor aimed at providing direct experimental information about the propellant ignition process. We further report on measurements of the ignition delays, and the propagation of the "ignition wave", performed using these sensors in experimental devices such as closed bombs and gun simulators.

In recent years, great strides have been made regarding the identification of the dominant loads that cause beam type vibrations of cannons during ballistic operation. These motions infringe upon shot accuracy because at projectile disengagement, the muzzle's kinematic state may compromise the projectile's intended flight path. This is particularly relevant to coil gun loading, which, unlike gas launch, applies the recoil load to the muzzle end of the barrel instead of the breech. This effectively places the barrel in a state of columnar compression during the launch cycle that may exacerbate structural vibrations known as gun "whip".

Through this work we show the development of small hybrid rocket thrusters intended for attitude control systems of small satellites and space-crafts. These thrusters have the advantage of being controllable and having higher specific impulse than the traditional solid fuel thrusters or hydrazine based thrusters. An internal ballistic model of the small hybrid thrusters is shown together with numerical results for the thrust-curves and other relevant parameters. The internal ballistic model is shown to be influenced by the regression speed and the need for higher regression speed fuel-oxidizer pairs is, hence, shown. In the present study we also show how to couple the hybrid combustion process with the unsteady flow. Moderately high specific impulses can be obtained by using polypropylene as fuel and NO2 as oxidizer. Detailed discussion is done on the construction of the thrusters as well as the engineering solutions for the combustion chamber, injector and nozzle. Several experimental thrust-curves are shown together with the estimation of total impulse and specific impulse.

Ignition transients are a critical and complex period of operation in solid rocket motors. During the ignition interval, many factors combine which must be taken into account in the design of the entire system. There are many methods in use to analyze different aspects of the ignition processes, and the level of analysis complexity varies with need and available resources. These methods vary from empirically based sizing rules to complex physics-based analyses. There are still many aspects of ignition transient behavior that have not been fully characterized, and research is continuing to fully understand all processes that occur during ignition transients in solid rocket motors. Keywords: ignition; solid rocket motor (SRM); propellant; ballistics; igniters; transient

The Sandia Hypervelocity Launcher (HVL) uses impact techniques on a two-stage light-gas gun to launch flier plates to velocities in excess of 10 km/s. An important problem in designing successful third stage techniques for impact launching fliers to such velocities is detailed understanding of the interior ballistic performance of the third stage. This is crucial for preventing melt and fracture of the flier plates during the extraordinary accelerations that they undergo (accelerations on the order of 10{sup 9} g are typical on the HVL). We seek to optimize HVL launch conditions in order to achieve two major goals: first, to maximize the potential launch velocity for a given flier, and second, to allow different flier configurations. One tool that we can apply in studying HVL performance is the use of multi-dimensional wave propagation codes. We have used such codes, particularly the Sandia Eulerian code CTH, to study a variety of interior ballistics issues related to gun performance and launcher development for almost ten years. Recently this work has culminated in a major contribution to HVL design, namely the capability to launch ``chunk`` fliers. `Me initial phases of design development were solely devoted to CTH computations that studied potential designs, identifiedmore » problems, and posed possible solutions for launching chunk fliers on the HVL. Our computations sufficiently narrowed the design space to the point that systematic experimental progress was possible. Our first experiment resulted in the successful launch of an intact 0.33 gram titanium alloy chunk flier to a velocity of 10.2 km/s. The thickness to diameter ratio of this flier was approximately 0.5.« less

Herein, we study the technique of improving the gun performance by installing a diaphragm in the pump tube of the gun. A CFD study is carried out for the 0.28 in. gun in the Hypervelocity Free Flight Radiation (HFF RAD) range at the NASA Ames Research Center. The normal, full-length pump tube is studied as well as two pump tubes of reduced length (approximately 75% and approximately 33% of the normal length). Significant improvements in performance are calculated to be gained for the reduced length pump tubes upon the addition of the diaphragm. These improvements are identified as reductions in maximum pressures in the pump tube and at the projectile base of approximately 20%, while maintaining the projectile muzzle velocity or as increases in muzzle velocity of approximately 0.5 km/sec while not increasing the maximum pressures in the gun. Also, it is found that both guns with reduced pump tube length (with diaphragms) could maintain the performance of gun with the full length pump tube without diaphragms, whereas the guns with reduced pump tube lengths without diaphragms could not. A five-shot experimental investigation of the pump tube diaphragm technique is carried out for the gun with a pump tube length of 75% normal. The CFD predictions of increased muzzle velocity are borne out by the experimental data. Modest, but useful muzzle velocity increases (2.5 - 6%) are obtained upon the installation of a diaphragm, compared to a benchmark shot without a diaphragm.

Radar signal processing approaches based on micro-Doppler signatures are of great potential for targets identification. This paper proposes a novel and perspicuous micro-motion model of Ballistic Missile(BM) targets. The correspondence between micro-Doppler(m-D) and micro-motions is analyzed. Finally, simulations demonstrate the proposition.

In the design and development of solid propellant rocket motors, the use of numerical tools able to predict the behavior of a given motor is particularly important in order to decrease the planning times and costs. This paper is devoted to present the results of the internal ballistics numerical simulation of the NAWC tactical motor n. 6, from ignition to burn-out, by means of a quasi-one-dimensional unsteady numerical simulation model, SPINBALL, coupled with a three-dimensional grain burnback model, GREG. In particular, the attention is focused on the effects on the SRM behavior of the erosive burning, total pressure drops and the cause of the pressure overpeak occurring during the last part of the ignition transient. The final objective is to develop an analysis/simulation capability of SRM internal ballistics for the entire combustion time with simplified physical models, in order to have reduced the computational costs, but ensuring an accuracy greater than the one usually given by zero-dimensional models. The results of the simulations indicate a very good agreement with the experimental data, as no attempt of submodels calibration is made, enforcing the ability of the proposed approach to predict the SRMs internal flow-field conditions. The numerical simulations show that NAWC n. 6 internal ballistics is completely led by the erosive burning, that is the root cause of the pressure peak occurring immediately after the SRM start-up.

In this paper the authors present their activity in the field of electromagnetic machine applications for aerospace solutions. A three stage electromagnetic accelerator is under construction to perform ballistic characterization of carbon-based nanocomposite materials for anti-debris application. Preliminary experiments as well as numerical simulation have been performed with promising results in terms of bullet's energy. Further implementation are needed in order to come closer the velocity of typical space debris (8km/s).

Interior ballistic model of spinning solid rocket motor (SRM) and a method of internal ballistic calculation were presented in this paper. The internal ballistic calculation equations were solved by fourth-order runge-kutta method. The calculation program and the law of pressure along with time were obtained. Through contrasting with test data, correctness of the theoretical model was verified. Finally, experimental methods and data disposal were discussed. The results indicate that the designed motor operates normally, the interior ballistic performance meets requirements, and the experiment can be done repeat. Furthermore, the experimental results have the guidance meaning for the pre-research design of spinning SRM.

A dual-density Rodman cone, with 1/d=7, has been found to be attractive for electromagnetic launch and may have adequate terminal ballistic performance. Stable flight is achieved from the correct distribution of mass within the projectile body. The design provides some flexibility for the armature in that all the armature mass is used for aerodynamic stability. Furthermore, the acceleration can be supported by a simple one-piece armature. A bore rider, which is less than 10% of the total mass, is needed for in-bore stability and structural support at the tungsten/aluminum interface. Work to date has focused on small-caliber applications, but substantial gains can be achieved when the bore size is increased to cannon caliber. General design principles for a mass stabilized projectile are presented. Nearly all aspects of launch, flight, and terminal effects as a function of bore size are addressed. Flight and terminal effects are computed from curve fits to existing experimental data. >

A railgun for the study of the interaction of meteorites with planetary surfaces and the interaction of space debris was developed. The authors have attained a velocity constantly higher than 5 km/s, with a maximum of 6 km/s. The importance of plasma confinement behind the projectile to prevent a leakage is shown to realize a high velocity. An effective acceleration using an aluminum rail is also shown. >

Proof-of-principle experiments of in-tube propulsion in which the propellant was supplied from the launch tube wall were conducted. The launch tube had a 25 mm × 25 mm square inner cross-section. A pair of opposing launch tube walls were made of polyacetal (POM), which was ablated using repetitive Transeversely-Excited Atmospheric (TEA) CO2 laser pulses. The laser beam was sent through a ZnSe window at the tube bottom to the symmetric parabolic surfaces of the projectile, and focused onto the ablator surfaces. The momentum coupling coefficient was measured from a vertical launch trajectory. When the breech of the launch tube was plugged, a momentum coupling coefficient of 4 mN-s/J at least was obtained. Even when the breech was connected to a vacuum chamber at a pressure lower than 20 Pa, the projectile could be vertically launched with a constant momentum coupling coefficient of 0.14 mN-s/J, the value of which was the same order as that obtained by directly irradiating a laser pulse onto a POM ablator.

To elucidate the spinning signatures of ballistic missile warhead, the model of spinning motion for ballistic missile warhead is established and the mathematical formulas for micro-Doppler signatures caused by spinning motion is derived. Then the time-varying micro-Doppler signatures are analysed with high-resolution time-frequency transform, and the micro-Doppler frequency is calculated based on time-frequency distribution. Finally the computer simulation confirms theoretical analysis, providing theoretical foundation for warheads recognition by micro-Doppler signatures.

Recent requirements for hypervelocity projectile launch for strategic U.S. Army missions have led to the proposal of a variety of gun propulsion systems. Among these systems are those that utilize solid propellant along with electrothermal-chemical (ETC) augmentation. Advanced solid propellant systems with complex grain geometry and loading configuration are also being investigated. In response to this need, the U.S. Army Research Laboratory (ARL) has, for a number of years, engaged in the study of solid propellant ignition using a hot, low molecular-weight plasma generated by ETC means. In addition, ARL is developing an ETC modeling capability using the next-generation interior ballistics code, NGEN, and a series of experimental studies in a ballistic simulator to generate code validation data. The current report demonstrates progress for application of the NGEN code to solid propellant, direct-fire gun systems in which various propellant configurations (i.e., grains and disks) are combined into a single charge in order to obtain higher-loading densities. The charge is ignited using the efflux from an ETC plasma capillary made of polyethylene. Comparison with pressure data measured in a 25-mm ballistics simulator, filled with disks of JA2 propellant as the main charge and a small amount of inert grains as filler, provides validation of the NGEN model with respect to the representation of the plasma efflux from an ablation capillary, into the chamber, and around propellant disks.

Recent requirements for hypervelocity projectile launch for strategic U.S. Army missions have led to the proposal of a variety of gun propulsion systems. Among these systems are those that utilize solid propellant along with electrothermal-chemical (ETC) augmentation. Advanced solid propellant systems with complex grain geometry and loading configuration are also being investigated. In response to this need, the U.S. Army Research Laboratory, in cooperation with Paul Gough Associates, is developing the Next Generation Interior Ballistics model, NGEN. The goal of this project is the development of an advanced interior ballistics model for use in the design and investigation of both indirect-and direct-fire gun propulsion systems. The current report demonstrates progress for application of the NGEN code to solid propellant, direct-fire gun systems in which various propellant configurations (i.e., grains and disks) are combined into a single charge in order to obtain higher-loading densities. The charge may be ignited using either a conventional (Benite) or an ETC plasma capillary. These notional charges are being investigated to determine code performance and assess the need for code upgrades, as well as to investigate propellant ignitability with Benite and plasma sources. Comparison with pressure data measured in a 25-mm ballistics simulator, filled with inert disks and grains, provides validation of the NGEN model with respect to the representation of the plasma efflux from an ablation capillary, into the chamber, and around propellant disks.

During the Columbia Shuttle Investigation, our office tried to identify a piece of on-orbit debris that originated from the Orbiter during its second day in space. This "Flight Day Two" (FD2) object was detected by UHF radar and tracked for 3 days before falling out of orbit. Extensive RCS measurements performed by the Air Force Research Laboratory (AFRL) and corresponding ballistic analysis by the USAF Space Command narrowed the potential candidates down to just two possible classes of objects: (1) a section of reinforced carbon-carbon (RCC) leading edge panel acreage; (2) a section of RCC "tee-seals". During the investigation, AFRL was asked to estimate the UHF RCS of various whole and fragmentary tee-seals to compare with the on-orbit UHF RCS observations. Since actual Orbiter tee-seal hardware, either whole or fractured, were not available, we predicted UHF RCS on various virtual tee-seal fragment geometries to confirm or eliminate the tee-seal as a candidate for the FD2 object. The paper summarizes our UHF RCS predictions, which definitively show that a whole or partial RCC tee-seal could not be the FD2 object. This left the RCC panel acreage as the only known object that fits both the on-orbit observed ballistic and UHF RCS data, a confirming piece of evidence in the Columbia investigation.

Hypervelocity impact experiments that reproduce on-orbit collisions between micrometeoroids or orbital debris and space structures are commonly performed by means of propellant-driven two-stage light-gas guns. Such devices accelerate projectiles using the thrust of a light propellant gas that is compressed to high pressure and temperature by a piston running in a pump tube. Though these guns have the unique capability of accelerating particles up to 9 km/s, many components of the gun must be checked and/or substituted after each shot making test sessions long and expensive. In order to have a lot of and many different types of hypervelocity impact data, the Center of Studies and activities for Space CISAS “G. Colombo” of Padua University developed a high-shot-frequency two-stage light-gas gun that can increase the shot repetition rate of standard facilities by a factor of 5 or more and at the same time reduce the shot cost by a factor of 2 or more. This is made possible through the use of special mechanical and diagnostic solutions that were designed to operate the gun for more than 50 shots in sequence without having to carry out maintenance operations. This article presents the design and operation of the CISAS two-stage light-gas gun damping system, which is one of the subsystems that makes it possible to achieve high-shot frequency. The damping system is in charge of controlling the piston oscillations in the pump tube, making it possible for the piston to withstand more than 100 shots without any damage. In particular, the damping system avoids piston strikes onto the gun head at the end of each compression stroke and allows the piston to be positioned at the base of the pump tube after each shot. The sensitivity of the piston oscillations to the damping operations and main subsystem design parameters were identified using numerical simulations, carried out according to a model that describes every working phase of the gun. Moreover, in this paper, the technical solutions for the damping system implementation are presented and the numerical predictions are compared with experimental results. For the CISAS high-shot-frequency gas gun, an efficient damping system proved to be a fundamental requirement to reliably accelerate 100 mg projectiles above 5 km/s and 70 mg projectiles at 5.5 km/s with a shot frequency of 10 shots per day at least, including the time needed for replacing the target and pumping down the target vacuum chamber.

Low-cost, micromachined, inertial sensors have been steadily emerging into the commercial marketplace. Some of these commercial off-the-shelf (COTS) sensors were evaluated for their insertion potential into military applications such as test and evaluation (T&E) and smart munition guidance. Performance requirements to provide time, space, position, and attitude information (TSPAI) and allow for navigation capabilities are fast approaching those needed, especially when integrated with a global positioning system (GPS) receiver. Artillery projectiles and rockets, instrumented with commercially available low-cost microelectromechanical systems (MEMS) accelerometers, spin sensors, and custom telemetry units, have been flight tested with good success. Test results suggest that many of these sensors are rugged enough for launch survivability and subsequent operation. Analysis of the accelerometer flight data shows good agreement with ground-based radar and other truth measurements. Ground tests of a MEMS gyroscope on a ballistic flight simulator have also been conducted. The results show promise for directly measuring a projectile's pitching and yawing behavior. This paper will present some of the test results and describe the challenges far using these devices as strapdown inertial measurement units (IMUs) on rolling projectiles.

The Army Research Laboratory has developed a scaleable, 3D, multiphase, computational fluid dynamics (CFD) code with application to gun propulsion (interior ballistics) modeling. The NGEN3 code, which incorporates general continuum equations along with auxiliary relations into a modular code structure, is readily transportable between computer architectures and is applicable to a wide variety of gun propulsion systems. Two such systems are the Army's Modular Artillery Charge System (MACS) and the Future Combat System (FCS). The MACS is being developed for indirect fire cannon on both current and developing (e.g., Crusader) systems. The efficiency of the MACS charge is dependent on proper flamespreading through the propellant modules; a process that has been repeatedly demonstrated in gun firings, successfully photographed using the ARL ballistics simulator, and numerically modeled using the NGEN3 code. The FCS requires weapons systems exhibiting increased range and accuracy. One of the technologies under investigation to achieve these goals is the electrothermal-chemical (ETC) propulsion concept, in which electrically generated plasma is injected into the gun chamber igniting the high-loadingdensity (HLD) solid propellant charge. NGEN3 code development and application to the MACS and FCS is currently a DoD HPC Challenge Project (No. 112) and is being greatly advanced by access to the DoD high performance computers (HPCs). Associate Fellow AIAA. Propulsion Physics Team Leader, Ballistics and Weapons Concepts Division, Weapons and Materials Research Directorate. Mechanical Engineer, Propulsion Physics Team, Ballistics and Weapons Concepts Division, Weapons and Materials Research Directorate. This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. INTRODUCTION A solid propellant gun system consists of a reaction chamber connected to a gun tube through which a projectile is guided once propelled by pressurization of the chamber. Chamber pressurization is accomplished by placing a solid propellant (SP) charge in the chamber and igniting it by various means. Current SP charges are generally complex structures consisting of hundreds or even thousands of distinct regularly formed (e.g., spherical, cylindrical) grains, which may be loaded in either regular or random arrangements. In addition to small-scale voidage between grains (i.e., porosity) many charges also contain large-scale voidage (i.e., ullage), which surrounds the entire charge (such as when the charge does not fill the entire chamber volume) or separates distinct subcharges (i.e., increments or modules) that together comprise the whole charge. The addition of energy to the chamber, usually near the gun breech, or rearmost end of the chamber, and in some cases through a tube extending along the centerline of the chamber, ignites the SP. In general, all of the grains are not ignited simultaneously, but an ignition flame spreads from the breech to the projectile base. The burning of the SP transforms chemical energy into heat as hot gases evolve from the surface of each grain of propellant. Initially the projectile resists movement allowing the pressure in the chamber to climb rapidly. Since the burn rate of the propellant is proportional to the pressure, hot gases are produced at an accelerated rate until peak pressure is reached in the chamber. Movement of the projectile down the gun tube, usually slight before peak pressure and much more significant afterwards, causes the chamber volume to increase, and generates rarefaction waves, which lower the pressure and thus the burn rate of the propellant. Upon ignition and burning, the gas dynamic flowfield in the gun chamber takes on a highly complex structure that includes the dynamics of propellant motion and combustion and various gas dynamic flow phenomena such as turbulent mixing, highly transient pressure waves, steep gradients in porosity and temperature, nonideal thermodynamics, and gas generation.

A nonlinear missile guidance controller with Thrust Vector Control (TVC) inputs for the interception of a theater ballistic missile is presented. In this paper, the midcourse phase of the missile is investigated, during which many non-ideal conditions must be considered, such as variation of the inertia of the missile and existence of the aerodynamic force, and an effect controller is designed to take into account the properties of TVC. Specifically, a 3 DOF optimal midcourse guidance law is designed to minimize the control effort and the distance between the missile and the target. The asymptotic stability of the overall system is proved by Lyapunov theory. Extensive simulations are conducted to validate the effectiveness of the proposed guidance law for TVC.

In a series of spectacular experiments conducted at the High Energy Laser Systems Test Facility (HELSTF), White Sands Missile Range (WSMR), NM, using 13- to 15-cm diameter, 40- to 60-g vehicles designed to fly on the 10 kW PLVTS pulsed carbon dioxide laser (1 kJ pulses for 30 microsecond duration at 10 Hz), Prof. Leik Myrabo of Rensselaer Polytechnic Institute (RPI) and Dr. Franklin Mead of the Air Force Research Laboratory's (AFRL) Propulsion Directorate, have been successfully flying laser propelled Lightcraft under a joint Air Force/NASA flight demonstration program. The axisymmetric Lightcraft vehicles are propelled by airbreathing, pulsed- detonation engines with an infinite fuel specific impulse. Impulse coupling coefficients have been measured with ballistic pendulums as well as a piezoelectric load cell and fall in the range of 100 to 200 N/MW. Horizontal wire-guided flights up to 400 ft, using a unique laser beam pointing and tracking guidance system, have demonstrated up to 2.0 G's acceleration measured by a photo-optic array. Spin-stabilized free-flights with active tracking/beam control have been accomplished to altitudes of 15.25 meters. This paper will summarize the progress made to date on the Lightcraft Technology Demonstration flight test program, since the first 12 - 14 July 1996, experiments at HELSTF.

An aerodynamic analysis of a pneumatic launch system capable of accelerating a 300 metric ton vehicle to 268 m/sec (600 mph) is presented. Vehicle acceleration is constrained to be less than or equal to 3-g (29.43 m/sec ). The launch system is a horizontal pneumatic ram (catapult) configuration. The ram is modeled as a cylindrical piston within a tube. Compressed air is injected at the upstream head end of the tube and the piston accelerated. The vehicle is attached to the ram and is accelerated with it. Injected air mass flow and total pressure are varied over time. The interior ballistics problem, fore and aft of the ram, is solved by using an unsteady, onedimensional method of characteristics technique. A genetic algorithm is used to find the optimum mass flow and total pressure vs. time at the injector, expressed as polynomials. A parametric study is then conducted on the performance of the system using the optimized polynomials. The results of the parametric study are compared to results obtained from a quasi-steady, Fanno flow model and to data obtained from a lab-scale model. Introduction The Highly Reusable Space Transportation (HRST) program was set up to to study concepts that would significantly lower launch costs. Part of this program includes the development of a ground-based launch system capable of providing first stage acceleration to the launch vehicle, resulting in a fuel mass reduction or a possible payload mass increase. A pneumatic launch system could offer a relatively simple and robust alternative to electromagnetic systems. The proposed system is a pneumatic ram consisting of a cylindrical piston in a circular tube, similar to the Navy's catapult launchers on aircraft carriers. The tube is assumed to be 7.92 m (26 ft.) in diameter and 1225 m (4019 ft.) in length. The length constraint comes from the kinematics of the ideal case of a constant 3-g acceleration (space shuttle maximum). The diameter constraint is more arbitrary but is representative of today's tunneling capabilities. Compressed air enters through a single injector at the base of the launch tube. The vehicle mass is assumed to be 300 metric tons. The piston mass is considered to be small in comparison to the vehicle mass and is neglected. 'Graduate Research Assistant, Aerospace Engineering; now Engineer/Scientist, Boeing 'Associate Professor, Aerospace Engineering; Senior Member AIAA * Director, Space Power Institute & Professor, Electrical Engineering Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. AH rights reserved. V«Mcl«

The use of fuel-rich solid (gas generator-type) propellants for hybrid propulsion affords some design and utilization efficiency advantages. Both forward and aft liquid injection control concepts are evaluated from the operational standpoints of ballistics, throttling, stability and extinguishment. Steady-state and non-steady ballistics analyses are employed for this evaluation. Stability of solid motor operation is enhanced by fluid injection with adequate injector pressure drop. Efficient throttling and reliable extinguishment are attained through a combination of solid propellant combustion tailoring, grain design, control valves and sensors. Initial results from a laboratory-scale slab combustor, combining a gas generator propellant with gaseous oxygen injection, are also presented.

A conceptual design of a very large aeroballistic range is presented, as are its operational characteristics and procedures. The proposed model launcher is a two-stage light-gas gun, having a launch tube diameter of 254 mm, and the capability of accelerating a 14 kg launch mass to 6.1 km/sec. The gun's 91.4 cm diameter piston is driven by pressurized helium. High pressures in the central breech are contained by a multiple disk arrangement. The blast tank and sabot separation tank are described, as are methods for arresting sabot segments. The conceptual design of the range itself includes a 3.3 m diameter test or flight chamber some 330 m in length. Provisions are made for testing of free flight models and tests in which the model is confined by a track system. Methods for model deceleration and recovery are described. Provisions required for future addition of advanced model launchers such as an electromagnetic launcher or ram accelerator are addressed. Siting and safety issues are also addressed.

A feasibility study was recently completed on a hopping transporter concept for locomotion on the moon. Termed a Lunar Pogo because it would operate similarly to a conventional pogo stick, this vehicle would accelerate up an inclined leg, pick up the leg, and then enter ballistic flight over the moon's surface. Upon recontacting the lunar surface, the Lunar Pogo would decelerate down the leg. Propulsion would be provided by expansion of gas against a piston. Operation would be partially conservative because much of the energy expended by the gas during takeoff would be recovered by compressing the gas during landing. Two models of the ballistics and propulsion have been set up to estimate performance. The simplified first-order model illuminates the thermodynamics of ideal operation. The second-order model, which includes realistic effects such as sliding, provides a better approximation to actual performance. In a one-man Lunar Pogo, an astronaut would typically make 15-m (50-ft) leaps at an average speed of 5-8 km (3-5 miles) per hour.

A program was initiated to design, build and evaluate the performance of a two-stage light gas gun, utilizing hydrogen gas, for use as an injector to an electromagnetic railgun. This effort is part of a project to develop a hypervelocity electromagnetic launcher for equation-of-state studies (HELEOS). The specific design features that enhance compatibility of the injector with the railgun, including a slip-joint between the injector launch tube and the coupling to the railgun, are described. The operational capabilites for using all major projectile velocity measuring techniques, such as in-bore pressure gauges, laser and CW (continuous wave) X-ray interrupt techniques, flash X-ray, and continuous in-bore velocity measurements using VISAR interferometry, are discussed. An internal ballistics code for optimizing gun performance has been utilized to interpret performance data of the gun. >

Millimeter wavelength radars are used to study plasma effects associated with ionized flow fields of projectiles launched at hypersonic speeds into a free-flight ballistic range. Two CW Doppler radars, at frequencies of 35 and 70 Gc/s, measure the nose-on backscattering radar cross sections during flight. The design and performance of the two radars are described in detail. A signal simulator provides absolute calibration. The purpose is to measure changes that occur in the radar cross sections of hypersonic projectiles caused by highly ionized flow fields. Under certain conditions the nose-on backscattering radar cross section of a blunt-nosed metal projectile decreases drastically when a thin, shock-produced layer of ionized gas covers the projectile. A theoretical analysis of this effect is given. Comparisons between theoretical predictions and experimental data show good correlation.

Studies were directed at the deflagration of heterogeneous, condensed media. Relations among ballistic properties were studied; generalized relations were derived and existing relations for testing the consistency of data from ballistic test motors were shown to be general to small error; errors in the recent literature were corrected. A new device for characterizing the ballistic properties of condensed media at high pressure with strands was devised and explored analytically. Self-pressurized constant pressure operation was shown along with capability to control the pressure level with a simple bang-bang control system. Special configurations to provide direct little difference measurements of ballistic sensitivities were presented. The Deur/Glick serial sandwich model for heterogeneous propellant combustion was modified to overcome the 'continuation' problem implicit in that model. Results showed ignition delays of correct magnitude and, with physical reasoning, demonstrated that the ZN methodology cannot be applied to heterogeneous propellants in its present form. Keywords: Solid rocket propellants, Solid propellant rocket engines.

The tests reported here were conducted to determine the effects produced on the ballistics of a liquid monopropellant round by advancing the point of venting of the primer within the propellant column. The tests were conducted with two case configurations in which the L/D ratios of the propellant columns were approximately 3.4 and 5.0. As the point of venting was advanced, muzzle velocity increased in rounds with the L/D ratio of 5.0, but decreased in those with the L/D ratio of 3.4. The amount of unburned propellant remaining in the case after firing increased in both case configurations. In both configurations, the pressure curves were transformed from curves with two distinct pressure peaks, through stages approximating plateau configurations, to curves with one well defined pressure maximum.

Closed-breech interior-ballistic equations were modified for predicting the performance of reactionless launchers. Test firings were made using two reactionless launchers of the Davis Gun type. Direct and high-low launchers were considered. The test firing data was compared with the curves obtained from an analog computer. Using the differences noted, certain factors of the original equations programmed into the computer were varied in order to bring the computer curves into close agreement with actual test results. These factors included heat loss coefficient, initial charge burnt, and in the case of the high-low gun the effective mean impetus of the propellant during the ballistic cycle. It was concluded that the knowledge gained with the computer can be used to optimize the performance of reactionless systems.