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A Hybrid Projectile (HP) currently under design at West Virginia University was simulated to estimate the effects of barrel launch angle and flight position of wing deployment. The projectile is similar to a standard 60mm mortar, except that is has been equipped to achieve extended range. A Simulink model was developed based upon external ballistics. The flight performance of the WVU-HP-60 was compared to a standard M720 60mm mortar. The developed HP was considered to be a tube-launched UAV, that transforms, not directly after launch but sometime after for optimal gliding, and must be modeled with different flight profiles because after transformation the aerodynamics drastically change. Two models of the UAV were created to allow for design of controllers. They were the launch model and the projectile flight model. It was found that the projectile may exit the barrel with a two degree variation of launch angle. The simulations show that range extension is still viable, with this barrel exit variation, to within 10% of the maximum achievable range. A confidence area was also developed to determine how far the launch angle and wing deployment position could stray and still maintain a significant amount of range extension.Copyright © 2012 by ASME

A concept of an aero-assisted pre-stage is proposed , which enables launch of both ballistic and aero-assisted launch vehicles from co nventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle so that their flight directions are coalign ed. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensu re prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet a irliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achi eved and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is acti vated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aeroassisted vehicles (spaceplanes) are considered.

A ballistic missile is one of the main targets of space target defense and often has an established trajectory. As the target, we are only interested in its trajectory. The intercepting missile, as a defense body, controls itself toward the target according to the target state observed information and the control strategies in real time. So, the kinematic equations, the dynamic equations and the control equations of the intercepting missile based on “transient equilibrium” hypotheses are more studied in detailed. In order to attack the target with the proportional guidance law in a three dimensional space, the dynamic equations of the proportional guidance law are established in two orthogonal planes which are projected onto by the motions of the missile and the target in the space. In the simulation phase, a distributed computer simulation system is established and the entire defense process is achieved in real time.

IllinoisRocstar LLC is enhancing and commercializing the computational capabilities of a comprehensive, integrated, massively parallel software suite for performing complex threedimensional multiphysics simulations. Referred to as the Rocstar Simulation Suite (RSS), its original target application was advanced solid rocket motor (SRM) modeling, analyzing internal multiphase flow ballistics, coupled combustion and internal grain deformation. The core fluid dynamics, solid mechanics, combustion, and thermal analysis component codes are based on research and development in LES turbulence modeling, multiphase flow, constitutive modeling, combustion chemistry, computational mechanics, thermal heat transfer, multicomponent coupling methodologies, uncertainty quantification, and multiscale modeling techniques. This paper will discuss the evolving architecture, why a multiphysics, multiscale modeling capability is important to the design and analysis of advanced solid propellant rockets, and presents a number of simulations that have recently been completed with Rocstar to illustrate the importance of these capabilities.

The U.S and the U.S.S.R. have sent seventeen successful atmospheric entry missions to Venus. Past missions to Venus have utilized rigid aeroshell systems for entry. This rigid aeroshell paradigm sets performance limitations since the size of the entry vehicle is constrained by the fairing diameter of the launch vehicle. This has limited ballistic coefficients (β) to well above 100 kg/m2 for the entry vehicles. In order to maximize the science payload and minimize the Thermal Protection System (TPS) mass, these missions have entered at very steep entry flight path angles (γ). Due to Venus' thick atmosphere and the steep-γ, high-β conditions, these entry vehicles have been exposed to very high heat flux, very high pressures and extreme decelerations (upwards of 100 g's). Deployable aeroshells avoid the launch vehicle fairing diameter constraint by expanding to a larger diameter after the launch. Due to the potentially larger wetted area, deployable aeroshells achieve lower ballistic coefficients (well below 100 kg/m2), and if they are flown at shallower flight path angles, the entry vehicle can access trajectories with far lower decelerations (∼50–60 g's), peak heat fluxes (∼400 W/cm2) and peak pressures. The structural and TPS mass of the shallow-γ, low-β deployables are lower than their steep-γ, high-β rigid aeroshell counterparts at larger diameters, contributing to lower areal densities and potentially higher payload mass fractions. For example, at large diameters, deployables may attain aeroshell areal densities of 10 kg/m2 as opposed to 50 kg/m2 for rigid aeroshells. However, the low-β, shallow-γ paradigm also raises issues, such as the possibility of skip-out during entry. The shallow-γ could also increase the landing footprint of the vehicle. Furthermore, the deployable entry systems may be flexible, so there could be fluid-structure interaction, especially in the high altitude, low-density regimes. The need for precision in guidance, navigation and control during entry also has to be better understood. This paper investigates some of the challenges facing the design of a shallow-γ, low-β entry system.

Current Mars entry, descent, and landing technology is near its performance limit and is unable to land payloads on the surface that exceed approximately 1 metric ton. One option for increasing landed payload mass capability is decreasing the entry vehicle's hypersonic ballistic coefficient. A lower ballistic coefficient vehicle decelerates higher in the atmosphere, providing additional timeline and altitude margin necessary for heavier payloads. This study analyzed the guided entry performance of concept low ballistic coefficient vehicles at Mars. A terminal point controller guidance algorithm was used to provide precision targeting capability. Accuracy at parachute deploy, peak deceleration, peak heat rate, and integrated heat load were assessed and compared to a traditional vehicle to determine the effects of lowering the vehicle ballistic coefficient on entry performance. Results from this study suggest that while accuracy at parachute deploy degrades with decreasing ballistic coefficient, accuracy and other performance metrics remain within reasonable bounds for ballistic coefficients as low as 1 kg/m2. As such, this investigation demonstrates that from a performance standpoint, guided entry vehicles with large diameters may be feasible at Mars.

The experiment reported in this article addresses whether polishing a rifle bore reduces the aerodynamic drag of bullets. It was hypothesized that by polishing a rifle bore the aerodynamic drag of bullets would be reduced because a smoother bore would result in smoother rifling marks and thus lower drag. A Remington 700 5R Mil-Spec chambered in 300 Winchester Magnum was used. The bullets used were a 155.5 grain Berger Fullbore Boat Tail and a 125 grain Nosler Ballistic Tip with near and far velocities measured with chronographs at 15 and 315 feet. G1 and G7 ballistic coefficients were calculated using the JBM online ballistics calculator including environmental conditions measured with a Kestrel 4500 Pocket Weather Tracker. These measurements did not support the hypothesis; in fact, the velocity losses were greater over the 300 foot chronograph separation and the ballistic coefficients were smaller after polishing the rifle bore. This suggests that polishing the rifle bore actually increased resulting aerodynamic drag on the bullets

Measuring an arrow's ballistic performance such as arrow velocity on impact, total time of flight and arrow shaft oscillation is challenging because of the dynamic nature of arrow flight. This challenge becomes increasingly difficult as the distance of the shot increases. It is also of great interest to bowhunters to understand the ballistic performance of arrows that include hunting broadheads. A miniaturized, sensory data acquisition system, located in the arrow tip and engineered to withstand the high accelerations experienced at launch and impact, enables the precise measurement of arrow ballistics in flight. By continuously recording arrow drag in flight, the system enables measurement of the ballistic performance of an arrow as it travels downrange. The authors have also built an adapter that is connected to the housing of the sensing system to allow for comparative ballistic tests to be performed on hunting broadheads. Here, we present results obtained using the sensing system to perform initial testing on two commercially available broadheads at shot distances of approximately 45 m.

The goal of many space missions to small bodies, such as the Hayabusa mission to Itokawa and the forthcoming OSIRIS-REx mission to asteroid 1999 RQ36, are to take samples from the surface of the body. The prospect of designing a transfer to an asteroid surface and choosing where on the surface to sample begs several questions related to the flow of material across an asteroid surface, possible regions of stable resting points, and the surface forces and slopes that a landing spacecraft will encounter. The detailed design of surface encounters rests heavily on the capabilities and design of the proposed spacecraft. A series of interesting case studies for the Hayabusa spacecraft and its successful touchdowns on the asteroid Itokawa surface were summarized in [81, 201, 64, 84]. In the following we take a more dynamics and mechanics oriented approach and describe the surface environment due to gravitational and rotational effects, describe measures of surface stability, and discuss some simple characterizations of the ballistics of bodies from the surface of an asteroid. This chapter does not probe the interesting question of how surface rovers should be designed and operated in order to achieve specific goals (see [85]).

During the flight of vehicles, their propulsion energy must overcome gravity, to ensure the displacement of air masses on vehicle trajectory, to cover both energy losses from the friction between a solid surface and the air and also the kinetic energy of reflected air masses due to the impact with the flying vehicle. The flight optimization by increasing speed and reducing fuel consumption has directed research in the aerodynamics field. The flying vehicles shapes obtained through studies in the wind tunnel provide the optimization of the impact with the air masses and the airflow along the vehicle. By energy balance studies for vehicles in flight, the author Ioan Rusu directed his research in reducing the energy lost at vehicle impact with air masses. In this respect as compared to classical solutions for building flight vehicles aerodynamic surfaces which reduce the impact and friction with air masses, Ioan Rusu has invented a device which he named free compression tube for rockets, registered with the ...

A study and a general model for the roll/pitch frequency lock-in phenomenonare made for long antitank kinetic energy projectiles and missiles in general to quantify the necessary projectile damage/defect required to cause the lock-in. Fin damage and body mass offset are modeled as the forcing forces and moments (asymmetries) causing the lock-in. Idealized Ž n damage is modeled.The corresponding pitch, side-slip, and roll equations are numerically integrated. The nonlinear induced roll moment is essential in coupling the pitching/yawing motions with that of the roll motion. A speciŽ c and complete projectile case study is presented, in which quantiŽ ed Ž n damage and rod mass offset are found to cause such lock-in. Computed  ight motion history (pitch, side slip, and roll) and the yaw ampliŽ cation factor provide insight into this lock-in behavior. The model is general and can be applied to any Ž nned projectile or missile.

A series of 2-D numerical simulations of the bulging of passive cassettes, penetrated by shaped charge jets, is described. The purpose of these simulations was to determine the factors affecting the speed of the bulging process in order to be able to design more efficient cassettes. Physical and geometrical parameters, of the different cassette constituents, were varied systematically. These included strengths, elastic moduli, thicknesses and densities of the materials involved.

Seeing that the space environment in which missiles flies has increasingly deteriorated, and the large threat has been worked by space debris, this paper builds a math model of the fall-point precision for impacting long-range strategic ballistic missile. By simulation and calculation, the influence of fall-point precision created by the impacting between missile and different sizes and different aspects space debris, and some conclusions can supply some references for warning of ballistic missiles' space fly.

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.

Ballistic missile tracking in time is one of the key technologies in the missile defence system. Aiming at problems about ballistic missile kinetic states real-time estimation, this paper investigated the method of ballistic missile tracking using dynamic model. Firstly, based on analyzing characteristics of ballistic missile's motion, this paper builds dynamics model of ballistic missile under the condition of gravity turn model; Secondly, based on analyzing parameter of dynamic model, this paper builds simplified kinetic model of ballistic missile; Lastly, this paper analyses how to chose kinetic states variable, and then educes ballistic missile kinetic states function and simulates the method of real-time estimation based on EKF, the result shows that the proposed method can track the ballistic missile, what is more, the estimation precision is high.

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.

Supersonic target cruise missile flying at zero is a dynamic practical simulator, which simulates the terminal flight character of anti-ship cruise missile. In order to simulator the flight height of anti-ship cruise missile, the target cruise missile must fly low. The transitional ballistic trajectory of supersonic target cruise missile from high to low is important for flying control. In this study, supersonic target cruise missile is investigated and the controllable particle math model is constituted in order to obtain the optimizing transitional ballistic trajectory. Principle of the minimum of optimum control is used to optimize the transitional ballistic trajectory. The results have an important theoretical value for fight control on transitional ballistic trajectory.

This paper provides an overview of the Antitank Missile Monte Carlo starting control points dispersion analysis conducted with the ballistic simulation. First, The disturbances that affect the flight of a missile are analyzed. Then a method of how to handle with the disturbances is presented to establish a ballistic simulation model with Monte carlo method, the steps of simulating the disturbances based on Matlab are also given. Taken account of the initial angular rates disturbance, rocket engine thrust misalignment, wind disturbance, the equation that described the dynamics of the missile mass center movement should be adjusted accordingly. In this paper, a common model is established to simulate the effect of disturbances. Results indicate that missile trajectory disturbances could be substantially reduced by lowering rocket engine thrust misalignment.

Ejecting interior ballistic equations of one submarine-launched cruise missile are established. And interior ballistic equations are simulated and computed. Basic changing rules of interior ballistic parameters are computed. Results indicate that gas-steam launch mode can be used to launch submarine-launched cruise missile. It is of some reference value for the application of gas-steam ejecting launch mode to the cruise missile.

In order to control the combustion stability of bulk-loaded liquid propellant gun, the expansion process of gas jet entering liquid was studied by high speed photographic system aiming at the stepped chamber and rectangular chamber. Based on the experiments, two-dimensional model of gas jet expansion process was developed. The unsteady interaction process of gas jet and liquid was simulated by the commercial software FLUENT. The influence of chamber structure on jet expansion was discussed. The calculated results approximately accord with the experimental results.

In this paper, trajectory correction of canard control anti-aircraft shells is researched. After launch, once targets make circumvent flight, we must adjust projectile control force to correct ballistic. Projectile free flying model and control flying model are established. Target Initial flying model and maneuvering flying model are assumed. According to the models of projectile and target, appropriate optimizing algorithm is selected, control force of projectile is optimizing calculated. We can draw the conclusion that the optimization algorithm is feasible.

This paper presents a novel fuel to be considered for hybrid rocket motors: solid methane. We are showing the potential that this fuel has for hybrid rocket motors from technical perspective as well as from economical perspective. We present an internal ballistic model used to design a hybrid rocket motor with solid methane. We also present a comparison table for solid methane and various cryogenic and non-cryogenic oxidizers. The maximum specific impulse is obtained for the solid methane-LOX pair with a value of ∼315 seconds under specific conditions. Although a solid methane-LOX motor shows higher than averages values for the regression speed, these values are still low by performance standards. We present two models that predict the regression speed increase for embedding wires and for oxidizer doping with a factor of ∼2–3. An experimental setup that can be used to verify these models is further presented in the paper. Finally, we present a potential application for solid methane hybrid rocket motors in the form of an upper stage propulsion unit for orbital/suborbital vehicles. We find a practical interdependence relation between structural optimization (dry mass) of the launcher and the performance parameters of the motor. We present a study case of such a vehicle intended to place a 50 kg payload on low earth orbit (300 to 500 km altitude). Developing such a launcher is both environmentally safer than current approaches and economically more efficient because methane reserves are a rich resource worldwide.

This report documents the theory and use of the BOOM computer program. BOOM is an exterior ballistics simulation program specifically designed to predict atmospheric flight of smart projectiles. The software contains a projectile dynamic model coupled to an open structure flight control system. The code can be run on PC, Unix, or Mac systems.

For the aerodynamic parameters identification of projectile and rocket under the influence of random wind, based on the control system sensitivity theory, the formation of the sensitivity and related mathematical models according to realistic systems are formulated. Sensitivity principle of the control system theory is applied to the identification of resistance coefficient. Analyzing missile system parameters, aerodynamic coefficients are identified using ballistic parameter sensitivity function. Resistance coefficient is identified through shooting observed coordinates data of ballistic. Trajectory data under the influence of random factors are re-calculated according to the identifying results to increase the identification error due to the inaccurate parameters. Through identification, the estimated mathematical model with the more precise ballistic parameters is established so as to enhance the accuracy of the launch of the missiles.

Reliable observation of space reentry vehicles is an issue which is worthy of concerning and discussing undoubtedly. The ballistic missile is a typical space reentry vehicle, and it flights along a destined trajectory with a high speed and may maneuver. The motion of the ballistic mission can be approximated by the interaction of the basic motion models that have the simpler forms, and can be observed through using ground radars and optical sensors. After the approximate motion models and observations are brought into the framework of UKF who aims at nonlinear estimation problems specially, the estimate track of the ballistic missile can be implemented. This paper describes the motion model of the ballistic missile firstly, and then the measurement equations of the different sensors and the basic model of the motion body are given out, and then the algorithms of interacting multiple model and UKF are expatiated on, and finally the estimate track is achieved in a simulation system and a conclusion is drawn.

The trajectory of incoming missile and shrapnel are analyzed to make sure that the incoming missile would be within the distribute range of bullets dispersed by the shrapnel. In the case of the center of bullet group hits the target, the miss distance is optimized with the angle of fire and direction angle as the optimization variables in MATLAB. An unconstrained optimization function ‘fminsearch’ is used. The optimal angle of fire and the direction angle of antimissile shrapnel are solved. The numerical calculation shows that the optimal angle of fire Θ∗ and direction angle φ∗ vary periodically with the orientation angle of the launching of shrapnel and the object needed to be protected. When the distance of the launching of shrapnel and the object needed to be protected increases from 60 to 1000 meters, θ∗ and φ∗ increase from 17.5°to 22.5°, from 67.4°to 79.7° respectively. When the muzzle velocity increases from 160 to 260m/s, and φ∗ change from 19.3°to 16.7°, from 48.5°to 54.4 ° respectively. The optimal angle of fire, the direction angle and their situation of change can be analyzed in the mathematical model for optimization of shrapnel gunnery.

Analysis of the interior ballistics is essential for the development of gun or propellant configurations. The granular solid propellants with high energy and fast burning rate produce a large thrust in extremely short time intervals. For the study of these, therefore, it is necessary of a numerical code for the two-phase flow of the interior ballistics. Recently, an interior ballistics code (IBcode) for the two-phase flow using the Eulerian-Lagrangian approach has been developed. The SIMPLE algorithm and the SMART scheme have been used for the IBcode. The ghost-cell extrapolation method has been used for the moving boundary with the projectile movement. In this study, a performance of the interior ballistics according to the position of the solid propellant in the chamber has been investigated using the IBcode. In previous researches, propellants had been evenly distributed in the chamber. In this study, however, three cases of the existence of empty space in the chamber at which the propellants are not evenly distributed have been considered; Propellants are located in the region near the base, propellants in the region near the breech, and propellants in the center of the chamber, respectively. The 7-perforated configuration of the solid propellant has been used in this research. The results have shown the performance variations of the interior ballistics according to solid propellant position in the chamber. The cases of the propellants located in the region near the base and breech have shown that the value of the negative differential pressure and the difference between the breech pressure and the base pressure are much higher than those of the propellants located in the center of the chamber. The case of the propellants in the center of the chamber is, therefore, more profitable to improve the performance of the interior ballistics.Copyright © 2011 by KSME

The U.S. Naval Air Warfare Center has participated in a program to develop an improved understanding of linear and nonlinear combustion instability in solid propellant rocket motors. One goal of this program was to develop a systematic database of motor and stability data. This paper describes the linear aspects of the motor e rings and analysis. The motors that were used had diameters of 127 mm and were 1.7 m long. The majority were loaded with an 88% solids reduced-smoke ammonium perchlorate propellant with a nominal burning rate of 6.1 mm/s at 6.9 MPa. In addition, motors have been e red that contain 1% 8- mm aluminum oxide, 90- mm aluminum oxide, and 3- mm zirconium carbide as stability additives in place of 1% ammonium perchlorate. Motor pressures ranged from 3.45 to 10.34 MPa and various grain geometries were tested. Pressure-coupled combustion response measurements were made at the nominal motor operating pressures. Motor performance and stability calculations were made using the Air Force Solid Performance Program and the Standard Stability Prediction Program for the motor cone gurations that were e red. The stability predictions were compared to the data obtained from the motor e rings. The results indicate that it is possible to theoretically predict linear motor stability for the class of motors discussed in this paper. It was also learned that higher-pressure motors tend to be less

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".

Este documento expone la metodología y los resultados obtenidos en el desarrollo de una investigación sistemática, que trata sobre a la implementación de misiones de cohetería experimental. En este proceso se obtuvieron, datos y resultados que permitieron corroborar de forma experimental, conceptos de propulsión, aerodinámica y balística propios de una aeronave tipo cohete sin elementos de control activo a bordo. Para el estudio e implementación de esta misión, se generó un método dividido en etapas por las cuales se guió el desarrollo del proyecto; primero se realizó la etapa de validación y eldiseño preliminar en forma teórica, se continuó con la etapa de simulación, utilizando un software especializado y culminó con la etapa del proceso constructivo y el procedimiento de disparo. La puesta en marcha de esta misión de cohetería experimental permitió alcanzar, propulsado por un combustible tipo “candy” compuesto por nitrato de potasio y sorbitol, un apogeo del cohete Ainkaa 1 de 780 m sobre la superficie de lanzamiento, con una velocidad de impacto de 526 Km/h, un tiempo de quemado de 3.2s entre otros datos, para configurar esta misión dentro de la categoría de alta potencia.

In order to estimate energy characteristics of gun propellant, which are hard to measure in experiments, an expert system for predicting the energy characteristic of gun propellant(ECPES) has been established based on a reasonable combustion model. The system has functions of energy characteristic estimation, knowledge management and formula optimization with characteristic constraints. ECPES is validated by experiments, and the result shows it could indicate energy characteristic of gun propellant, with an error less than 3%.

With regard to the research object of a kind of rocket projectile, the influence of gun muzzle vibration by jet impact is considered. Both the launching dynamic model and the intensive degree model are founded. The circumstances of firing time partition for 0.3s to 0.5s are simulated numerically. The results are compared, and the way to improve the intensive degree is presented as well.

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).

A fire fighting cannon is fire-fighting equipment developed for fighting against forest fires; it is considered the tree crown fire and the inaccessible areas which are dangerous for man to get close as extinguishing object. The forest fire fighting cannon is launched by pneumatic method; the timing device within the extinguishing shell will be triggered after launching. After the free-flying and the parachute deceleration phase, the shell will detonate while the temperature in the fire field detected by the temperature sensor arrives at the setting value. The control system of the fire extinguishing cannon was designed firstly, the fly ballistic trajectories were analyzed and calculated, which are involved with the interior ballistic in the process of launching, the exterior ballistics analysis in the state of free-flying with non-parachute and the process of deceleration flight with parachute open. The ballistic calculation provides control with the necessary parameters.

An interior ballistic simulation of the canister launched missiles has been developed with the help of FLUENT. One of the special features of this simulation is the almost uniform acceleration which occurs 0.25s after startup and keeps the missile shockless during ejection. The calculation was done efficiently in FLUENT after setting all the required parameters. Several figures have been given to show the velocity and pressure of the flow field in the reservoir. The variations of the velocity and acceleration with time show that the simulation and the test data of the experiments are in good agreement with each other. This model can be used to analyze the similar launch procedures.

For hybrid rocket engines, there are some unique relationships between the parameters, such as the fuel regression rate, oxidizer mass flow rate, equivalence ratio and so on, which determine the burning properties. These relationships are strongly related to the engine performance. Our final goal of this study is to establish the optimum design method which considers the engine performance of hybrid rocket engines for practical hybrid rockets varying with time. In this paper, we focused on the time variation of ballistic parameters. First, we established the calculation method which well simulates the burning condition of the swirling-oxidizer-flow-type hybrid rocket engine. Next, we predicted the time variation of the burning properties and ballistics parameters using this calculation method and some experimental burning data. The simulation results showed little difference in the time variations of the combustion pressure and thrust from the experiment. The simulated values of the engine efficiencies of Isp, C* and CF obtained well agreed with those experimentally. The simulation method proposed here is appropriate for predication of time variation of ballistic parameters and may be useful as a design tool for hybrid rocket engines.

The problem of evaluating the performance of an uncertain guidance loop system is considered, when the uncertainty is described in terms of an integral quadratic constraint. The idea of the approach proposed in this paper is to determine the set of all possible state vector values at the end of the flight that are compatible with the model uncertainty and determine the upper and lower bound for the miss distance by solving two optimization problems on this set. The main theoretical result of the paper shows that this is a quadric that can be determined by the solving a differential Riccati equation and a set of coupled linear differential equations. Unfortunately, numerical experiments showed that the method is practically unusable for realistic models of guidance loops. A single integral quadratic constraint is a poor representation of the usual model uncertainties, meaning that the results are certainly too conservative. In addition, the optimization problem turns out to be badly conditioned in all cases of interest for homing guidance loop analysis. This is illustrated using a contrived numerical example based on a relatively well-behaved linear time-invariant system. Although in its present form, the approach taken in this paper did not achieve the results that were hoped for, the ideas advanced here, and some of the technical results are interesting in themselves and may be useful in future research on this theme.

To solve problems of low efficiency, long period and not handling coupling in solid rocket motor traditional design, multidisciplinary design optimization method of solid rocket motor was specified through comparison of typical MDO method and multi-objective collaborative optimization model of solid rocket motor MDO was constructed. Aiming at computation difficulties of CO, paper put forward some approaches to solve the problem. Finally, to reduce solid rocket motor computation expense, three most important disciplinary approximate analysis model of shell, thermodynamic calculation and interior ballistics model were formulated. This research will be valuable for solid rocket motor design using MDO theory.

This chapter is devoted to the discussion of simulation results of several space flight scenarios ranging from movements in Earth or planetary orbits via aerobraking and aerocapturing missions up to re-entry problems, and also of artillery ballistics.

The problem of evaluating the performance of an uncertain guidance loop system is considered, when the uncertainty is described in terms of an integral quadratic constraint. The idea of the approach proposed in this paper is to determine the set of all possible state vector values at the end of the flight that are compatible with the model uncertainty and determine the upper and lower bound for the miss distance by solving two optimization problems on this set. The main theoretical result of the paper shows that this is a quadric that can be determined by the solving a differential Riccati equation and a set of coupled linear differential equations. Unfortunately, numerical experiments showed that the method is practically unusable for realistic models of guidance loops. A single integral quadratic constraint is a poor representation of the usual model uncertainties, meaning that the results are certainly too conservative. In addition, the optimization problem turns out to be badly conditioned in all cases of interest for homing guidance loop analysis. This is illustrated using a contrived numerical example based on a relatively well-behaved linear time-invariant system. Although in its present form, the approach taken in this paper did not achieve the results that were hoped for, the ideas advanced here, and some of the technical results are interesting in themselves and may be useful in future research on this theme.

A concept for propelling a projectile to hypervelociti es is described, and the feasibility is demonstrated by numerical simulations. In theory, the concept employs imploding blast waves to accelerate a projectile as it travels down a launch tube. The launch tube can be open to the atmosphere or sealed and maintained at some low pressure to minimize internal drag. The launch tube has a liner that contains a suitable explosive or energetic material. The explosive is configured with inert annular rings to prevent upstream detonation. A suitable trigger detonates each explosive ring sequentially as the projectile passes. The resulting blast wave causes an elevated pressure on the afterbody of the projectile. The acceleration continues until the projectile exits the launch tube.

Ballistic impact experiments were performed on ceramic laminate targets at three scale sizes, nominally 1 3 , 1 6 , and 1 12 , to quantify the effects of scale on various responses, in particular, the ballistic limit velocity. The experiments were carefully designed and controlled so that the different scale sizes were high fidelity replicas of each other. A variety of responses, such as residual projectile quantities, hole size, and the extent of bulging, were measured. Some of the measured quantities showed little or no dependence on scale size, whereas other quantities, particularly the ballistic limit velocity, were found to vary with scale size. The percentage difference was quantified, and the results extrapolated to estimate full-scale response from the subscale tests.

A new way to attack the aerial target has been put forward, while taking use of the height and velocity dominance of the high dynamic Near Space aircraft, which is a new operational mode. The feasibility of Near Space un-powered guided bomb when attacking aerial target is described mathematically. Then the precise ballistic model of the guided bomb is established. The Response Surface Models (RSM) of the average velocity and final velocity of the guided bomb are constructed based on Uniform Design experimentation method. The RSM is optimized with Genetic Algorithm based on grade penalty function. A set of feasible solutions are obtained through the simulation example.

An efficient method was needed to perform ballistic testing using the 0.22-cal. fragment-simulating projectiles (FSP) at low subsonic velocities in order to evaluate very thin lightweight composite and metallic materials at around 1 lb/ft2 areal density. This technical note outlines the custom gun system developed to enable this ballistic testing.

Base on the launch mechanism and structure characteristic of rarefaction wave gun, interior ballistic gas dynamic model of it applying the inertial breech that considering the dynamic opening process of went equipment is established. By calculating the model the distribution of interior ballistic parameters in the launch process are also discussed, and then by contrasting the distribution of press, temperature, recoil force and recoil momentum of the same type general gun. It is concluded that rarefaction wave gun can extremely reduce recoil momentum to 88.3% and reduce the barrel temperature to 34.1% without effect the muzzle velocity under the same loading condition.

The paper purpose is to dignify some aspects regarding the calculus model and technical solutions for small multistage suborbital launchers used to test spatial equipment and scientific measurements. The calculus methodology consists in numerical simulation of launcher evolution for different start conditions. The rocket model presented will be with six DOF and variable mass. The results analyzed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize small suborbital multistage launchers recycling military rocket engines.