Projectile Configurations Research Articles

Control Surface Geometry Surrogate-Based Optimization for Spin-Stabilized Projectile Course Correction

In the present paper, a surrogate-based methodology is developed and applied to the design of a control surface placed on an artillery projectile. An artificial neural network is used to estimate the aerodynamic contribution of the control device from numerical simulations. These estimations are coupled to a flight mechanics code to evaluate the trajectories of controlled projectiles. From these six-/seven-degree-of-freedom computations, the trajectories modifications are modeled using kriging. A correction device should minimize the distance to a ellipse (with as the projectile standard deviation of the projectile dispersion) around the mean impact point while ensuring problem-specific constraint fulfillment. The kriging databases are sequentially enriched with the impact point position of the projectile configurations maximizing the expected improvement criterion until a convergence state, assessed by leave-one-out cross validation, is reached. Computational-fluid-dynamics-based evaluations of the optimum configuration coefficients provide a refinement of the neural network databases in the relevant area of the design space. A course correction corresponding to a decrease of two standard deviations in the lateral direction and an increase of two standard deviations in range is achieved, respectively, by reducing the pitching moment of the projectile and increasing its lift-to-drag ratio.

Impact point prediction guidance based on iterative process for dual-spin projectile with fixed canards

The fixed canards configuration of a dual-spin projectile makes it difficult to apply the traditional guidance law. In this study, a modified impact point prediction guidance strategy based on an iterative process was developed for a class of dual-spin projectiles with fixed canards, to reduce the impact point dispersion. The guidance strategy is dependent on the modified projectile linear theory to rapidly predict the flight states and the impact point. For projectiles with control applied to the trajectory, the modified projectile linear theory method is known to achieve poor impact point prediction. To improve the prediction accuracy, improvements were made to the modified projectile linear theory by considering the products of the yaw rate and other small quantities. The guidance strategy is based on the iterative process for the continuous adjustment of the expected output of the roll angle of the course correction fuze, to minimize the direction error between the predicted impact point and target location. Studies were conducted on a model dual-spin projectile configuration to demonstrate the guidance details. The numerical simulations indicate that the proposed guidance strategy can effectively reduce the projectile impact point dispersion.

Smart Projectile Parameter Estimation Using Meta-Optimization

Parameter estimation and system identification of smart projectiles is an important and commonly used industrial tool. Existing methods rely on good initial estimates of the parameters to avoid local minima and ensure proper convergence. New projectile configurations may include highly nonlinear dynamics or unknown control parameters that cannot be known a priori. A new method for projectile parameter estimation is proposed that combines an output error parameter estimation algorithm with meta-optimization. Meta-optimization uses a suite of optimizers in an intelligent manner to reliably minimize a cost function. This new method is applied to the identification of a smart projectile system equipped with microspoilers using simulated spark range data. The method is able to reliably estimate the aerodynamic coefficients of the projectile body as well as the properties of the control mechanism based on a fit of multiple trajectories.

Debris dispersion analysis for the determination of impact conditions via traceback technology

A method for predicting the initial impact conditions including projectile shape and impact velocity is presented through analysis of the dispersion characteristics of debris by means of a traceback methodology. For the determination of the impact conditions, the diameter of the penetration hole in the target, and the dispersion area of the debris on the witness plate were investigated. Based on the results, the projectile configuration defined by the nose angle (NA) and the length to diameter ratio (L/D) together with the initial impact velocity was obtained. NA is associated with the penetration diameter DP and also debris dispersion diameter D50. D50 means 50% containing debris of all dispersions from the center of witness structure. Furthermore, L/D is related to DP and D95, in which D95 is the dispersion diameter containing 95% debris. The projectile diameter can be determined by the relationship between the projectile shape (NA and L/D) and debris dispersion (D50 and D95). Meanwhile, the shape factors identified from the contact surface between the projectile and the structure were designed to modify the de Marré’s ballistic limit equation. From the results, it was found that DP increased linearly with increasing impact velocity, and the slopes of DP against impact velocity also increased as NA increase. The initial velocity has been then finally determined by the DP-impact velocity slope and the modified de Marré’s equation. Based on the methodology, the shape and initial impact velocity of the projectile could be predicted within 5% error.

Ion-induced secondary electron emission from metal surfaces

Using a helium ion hitting various metal surfaces as a model system, we describe a general quantum-kinetic approach for calculating ion-induced secondary electron emission spectra at impact energies where the emission is driven by the internal potential energy of the ion. It is based on an effective model of the Anderson−Newns-type for the subset of electronic states of the ion−surface system most strongly affected by the collision. Central to our approach is a pseudo-particle representation for the electronic configurations of the projectile which enables us, by combining it with two additional auxiliary bosons, to describe in a single Hamiltonian emission channels involving electronic configurations with different internal potential energies. It is thus possible to treat Auger neutralization of the ion on an equal footing with Auger de-excitation of temporarily formed radicals and/or negative ions. From the Dyson equations for the projectile propagators and an approximate evaluation of the self-energies, rate equations are obtained for the probabilities with which the projectile configurations occur and an electron is emitted in the course of the collision. Encouraging numerical results, especially for the helium−tungsten system, indicate the potential of our approach.

Development of the ILR-33 “Amber” sounding rocket for microgravity experimentation

This paper gives an overview of the development of the ILR-33 ”Amber” sounding rocket designated for microgravity experiments, that is under development at Institute of Aviation in Warsaw, Poland. The lack of an easily accessible and affordable platform for this kind of research was one of the key reasons for this work. The proposed design enables performing experiments in microgravity for almost 150 seconds with an apogee over 100 km. Combining these results with a relatively low price per launch and short deployment time gives a possibility to establish a firm position on the dynamic market. This article describes also the rocket structure and the vehicle's capabilities. The proposed design utilizes a hybrid rocket motor with High Test Peroxide as an oxidizer along with two reusable solid rocket boosters. The early phase analysis of the rocket configuration and propellant considerations are also presented in the paper. Furthermore, there have been already several on-ground test performed such as: wind tunnel research and motor firings. The proposed design is considered as an introduction to small launch vehicle technology.

Perforation mechanics of 2024 aluminium protective plates subjected to impact by different nose shapes of projectiles

This paper focuses on the mechanical behaviour of aluminium alloy 2024-T351 under impact loading. This study has been carried out combining experimental and numerical techniques. Firstly, experimental impact tests were conducted on plates of 4mm of thickness covering impact velocities from 50m/s to 200m/s and varying the stress state through the projectile nose shape: conical, hemispherical and blunt. The mechanisms behind the perforation process were studied depending on the projectile configuration used by analyzing the associated failure modes and post-mortem deflection. Secondly, a numerical study of the mechanical behaviour of aluminium alloy 2024-T351 under impact loading was conducted. To this end, a three-dimensional model was developed in the finite element solver ABAQUS/Explicit. This model combines Lagrangian elements with Smoothed Particle Hydrodynamics (SPH) elements. A good correlation was obtained between numerical and experimental results in terms of residual and ballistic limit velocities.

Effect of firing conditions & release height on terminal performance of submunitions and conditions for optimum height of release

Abstract Submunitions should exhibit optimum terminal performance at target end when released from certain pre-determined height. Selection of an optimum height of release of the submunitions depends on the terminal parameters like forward throw, remaining velocity, impact angle and flight time. In this paper, the effects of initial firing conditions and height of release on terminal performance of submunitions discussed in detail. For different height of release, the relation between range and forward throw is also established & validated for a number of firing altitude and rocket configurations.

Conception and Manufacturing of a Projectile-Drone Hybrid System

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.

Influence of impact conditions on plasma generation during hypervelocity impact by aluminum projectile

For describing hypervelocity impact (relative low-speed as related to space debris and much lower than travelling speed of meteoroids) phenomenon associated with plasma generation, a self-developed 3D code was advanced to numerically simulate projectiles impacting on a rigid wall. The numerical results were combined with a new ionization model which was developed in an early study to calculate the ionized materials during the impact. The calculated results of ionization were compared with the empirical formulas concluded by experiments in references and a good agreement was obtained. Then based on the reliable 3D numerical code, a series of impacts with different projectile configurations were simulated to investigate the influence of impact conditions on hypervelocity impact generated plasma. It was found that the form of empirical formula needed to be modified. A new empirical formula with a critical impact velocity was advanced to describe the velocity dependence of plasma generation and the parameters of the modified formula were ensured by the comparison between the numerical predictions and the empirical formulas. For different projectile configurations, the changes of plasma charges with time are different but the integrals of charges on time almost stayed in the same level.

Feasibility of a low-cost sounding rockoon platform

This paper presents the results of analyses and simulations for the design of a small sounding platform, dedicated to conducting scientific atmospheric research and capable of reaching the von Kármán line by means of a rocket launched from it. While recent private initiatives have opted for the air launch concept to send small payloads to Low Earth Orbit, several historical projects considered the use of balloons as the first stage of orbital and suborbital platforms, known as rockoons. Both of these approaches enable the minimization of drag losses. This paper addresses the issue of utilizing stratospheric balloons as launch platforms to conduct sub-orbital rocket flights. Research and simulations have been conducted to demonstrate these capabilities and feasibility. A small sounding solid propulsion rocket using commercially-off-the-shelf hardware is proposed. Its configuration and design are analyzed with special attention given to the propulsion system and its possible mission-orientated optimization. The cost effectiveness of this approach is discussed. Performance calculation outcomes are shown. Additionally, sensitivity study results for different design parameters are given. Minimum mass rocket configurations for various payload requirements are presented. The ultimate aim is to enhance low-cost experimentation maintaining high mobility of the system and simplicity of operations. An easier and more affordable access to a space-like environment can be achieved with this system, thus allowing for widespread outreach of space science and technology knowledge. This project is based on earlier experience of the authors in LEEM Association of the Technical University of Madrid and the Polish Small Sounding Rocket Program developed at the Institute of Aviation and Warsaw University of Technology in Poland.

Aerodynamic characteristics of a wrap-around fin rocket

Purpose – The purpose of this paper is to investigate the aerodynamics characteristics (especially the side force/moment and rolling characteristics), to analyze the impacts generated by different parameters of wrap-around fins (WAFs) and to find the corresponding mechanism. Design/methodology/approach – The paper has adopted three different types of WAFs for the rocket configurations and the sub-regions divided technology to investigate the lateral and rolling characteristics of WAFs, including the fins with variations in span to chord ratio, thickness, leading-edge sweep, curvature radius, fin numbers, setting angles and rotated angles. Simulations have been performed at Mach numbers from 3 to 4 through an angle-of-attack range of about 0° to 10° and at model rolling angles of 45° to 90°. Findings – The paper shows that the WAF configurations can greatly improve the longitudinal stability and enhance the longitudinal aerodynamic characteristics for the whole rocket. The total drag of the whole rocket is...

Aerothermal Postflight Analysis of the Sharp Edge Flight Experiment-II

The aerothermal instrumentation of the Sharp Edge Hypersonic Flight Experiment SHEFEX-II, which was launched from Andøya Rocket Range on top of a two-stage rocket configuration, provided very useful flight data. Complementary to the pressure data, heat flux rate and surface temperature were measured at selected locations on the payload along the complete ascent and descent trajectory. Measured heat flux rate and pressure fluctuations show excellent correlation with angle-of-attack variations. Calculated heat flux evolution using a hypersonic boundary-layer code along the trajectory is in reasonable agreement with measured heat flux data during flight. Boundary-layer transition was measured at a boundary-layer edge Reynolds number of approximately for both ascent and descent phases. The complementary analytical and numerical study for the flight rebuilding at selected flight points provided very useful data and allowed better understanding of the physical phenomena. Finally, a combined analytical–numerical approach was validated and applied for the computation of the heat flux rate along the flight trajectory. This verified method can be used for the design and analysis of future hypersonic flights of slender vehicles.

Computational Fluid Dynamics Simulation of Hybrid Rockets of Different Scales

CFX software is used to simulate different hybrid rocket configurations, applying liquid as the oxidizer and paraffin as the fuel. This work is the prosecution of a previous paper analyzing liquid injection in a lab-scale hybrid rocket. It is focused on the formulation of the most suitable simulation technique to represent another type of liquid injector, compared with the one described in the previous paper. It also aims at extending the computational fluid dynamics simulation approach to hybrid rockets of larger scales. To validate computational fluid dynamics output, experimental results coming from both a laboratory scale and an increased-scale engine have been used. The different geometries studied include an increased-scale engine with a cylindrical grain having no diaphragm, the same rocket with a one-hole diaphragm inside the fuel grain, and a lab-scale rocket with a one-hole diaphragm. Simulations are steady state, and combustion derives from a single-phase chemical reaction. Liquid injection is fully simulated for the oxidizer, but paraffin entrainment is neglected for the fuel. Computational fluid dynamics results show a good agreement with the corresponding experiments for the ballistic parameters of interest in this study: chamber pressure, efficiency, and . Computational fluid dynamics results prove that the simulation technique proposed can be applied to any hybrid rocket scale.

Parametric study of different fins for a rocket at supersonic flow

Various parameters including the fins with variable span to chord ratio, curvature radius, and setting angle have been investigated between the flat fin and wrap around fin (WAF) rocket configurations at supersonic flow. The results show that under the same flight condition, the flat fins can provide a higher lift and pitching moments than the WAFs. Due to the symmetric effect, any extra side forces, moments as well as the self-induced rolling characteristics will be not generated as compared to the WAF configurations. The WAFs can greatly improve the longitudinal stability and enhance the longitudinal aerodynamic characteristics for the whole rocket. The static pressure distributions at different chordwise positions together with the force variations around the fins have been obtained computationally and analyzed in detail.

Single-Shot Performance of RAFIRA

The rapid fire railgun (RAFIRA) developed at the French-German Research Institute of Saint-Louis (ISL) is a medium-caliber (25 mm x 25 mm) system designed to obtain very high fire rates (>50 Hz). It has an acceleration length of 3 m and is therefore comparable with existing close-in weapon systems such as Goalkeeper or Phalanx. In this paper, the single-shot performance of RAFIRA is presented. While RAFIRA has, in the past, mainly been used for multishot studies, its 3.4-MJ power supply allowed testing speed limits of the brush armature technology used at the ISL. The combination of light-weight sabots made from glass fiber-reinforced plastic and multiple metal fiber bundles acting as armatures allowed muzzle velocities in excess of 2 km/s while exhibiting excellent sliding contact performance throughout the launch. Corresponding experimental findings for varying projectile configurations and primary energies are presented and discussed.

Numerical Simulation of Hybrid Rockets Liquid Injection and Comparison with Experiments

In this paper, CFX is used to simulate different hybrid rocket configurations applying liquid as the oxidizer and paraffin wax as the fuel. This work is intended as the prosecution of a previous study about hybrid rockets with diaphragms of different geometries inside the combustion chamber, where was injected in a gaseous phase. In this work, liquid injection is introduced, together with droplets vaporization and their coupling with the Eulerian gas phase, in terms of both heat and momentum exchange. The main objective is the description of the numerical models to be applied when liquid is injected. To validate computational fluid dynamics output, experimental results coming from a laboratory scale engine have been used. The different geometries studied include an engine with a cylindrical grain having no diaphragm and the same rocket with a four-hole diaphragm at 24% of the grain length. The simulations are steady state, and combustion derives from a single-phase chemical reaction. Liquid injection is fully simulated for the oxidizer, but paraffin entrainment is neglected for the fuel. Computational fluid dynamics results show a good agreement with the corresponding experiments, concerning the ballistic parameters of interest in this study: chamber pressure, efficiency, and characteristic velocity . Computational fluid dynamics predicts (both for gas and liquid injection) a higher efficiency for the rocket geometries provided with a diaphragm compared to the same geometries without a mixing device, and this is in accord with experiments.

Experimental Investigation of an Alternative Rocket Configuration for Rocket-Based Combined Cycle Engines

Experimental tests are conducted to evaluate two rocket exhaust configurations within an Rocket-Based Combined Cycle engine model. Each configuration uses a single rocket chamber but expands the rocket exhaust to a different geometry. One configuration uses two traditional circular rocket nozzles whereas the other uses two annular exhaust geometries based on the Exchange Inlet design. Wall pressure data are collected along the length of each configuration to evaluate the air entrainment. The annular configuration is shown to increase the entrainment ratio by as much as 28% over the traditional circular rocket design. The annular configuration is also shown to choke the incoming airflow at a lower rocket chamber pressure more uniformly across the engine cross section. Total pressure measurements are also obtained at the engine exit plane, which show that the annular configuration yields an average Mach number 57% higher than the circular configuration under conditions close to those for maximum air entrainment.

Periodic projectile linear theory for aerodynamically asymmetric projectiles

A new analytical tool is proposed to aid in the design and performance evaluation of advanced guided projectile concepts. A projectile linear theory applicable to aerodynamically asymmetric configurations is created, leading to a linear, periodic dynamic system. Utilizing concepts in Floquet theory, stability of asymmetric projectile configurations is explored. While stability of many asymmetric projectile configurations can be accurately predicted using averaged linear, constant coefficient system dynamics, there are some configurations where the use of linear periodic systems theory is required for accurate stability prediction. This fact is shown by comparing stability boundaries of an example projectile configuration using the conventional projectile linear theory model and the new periodic projectile linear theory model.

Sharp Edge Flight Experiment-II Instrumentation Challenges and Selected Flight Data

The Sharp Edge Flight Experiment II was launched from the Andoya rocket range in Norway on 22 June 2012, consisting of an extensively instrumented scientific payload on top of a two stage rocket configuration. With an apogee of about 177 km, the vehicle achieved flight velocities up to corresponding to Mach numbers up to 9.3. More than 96% of the pressure sensors provided excellent data during the ascent and descent phases. Measured pressure data from sensors at different locations of the scientific payload show consistent results concerning the aerodynamic behavior of the vehicle along the complete trajectory. Pressure fluctuations measured during flight show an excellent correlation to angle-of-attack variations. Calculated pressure coefficients from a computational fluid dynamics analysis at selected trajectory points are in good agreement with the measured pressure data.