Missile Wing Research Articles

A numerical analysis of pressure characteristics in the deflector jet pilot stage valve with an innovative deflector deflection

Electric-hydraulic servo-valves are extensively used in the aerospace, mechanical, and aerodynamics industries to control missile wings or airplanes precisely. Aiming to reduce undesired flow phenomena inside the deflector jet pilot valve, innovative deflections of deflector are proposed in this paper. A numerical analysis is performed using STAR CCM + to investigate the pressure and flow characteristics inside the pilot stage of deflector jet servo-valve. For observing the pressure, cavitation and flow characteristics, the Eulerian multi-phase model, VOF phase interaction, and realizable k−ε model for the turbulent flow are used during numerical simulations. The supply and outlet pressure are kept at 24 MPa and 1 MPa, respectively. To verify the calculated numerical results for turbulent flows, the velocity distribution to the V-groove exit of the deflector jet pilot stage is calculated and discussed with the theoretical results. The numerically calculated results show good agreement with published analytical results. The result shows that by the innovative rotation of the deflector both in clockwise and anti-clockwise directions, significant pressure in the receivers is obtained than traditionally moving deflector to the left and right for the same supply pressure. Moreover, when the deflector is positioned clockwise or anti-clockwise by 5°, less cavitation is observed than in the traditional model. The key finding of this paper shows that innovative structural optimization of the amplifier disk and deflector is of utmost importance.

Static and dynamic characteristics of supersonic cruise missile with damaged wing

AbstractAccurately evaluating the aerodynamic performance of the missile with damaged structures is very important for the subsequent flight control strategy. At present, few researchers have studied the aerodynamic characteristics of damaged supersonic cruise missiles. Based on CFD (computational fluid dynamics) solutions and the dynamic derivative identification method, the differences in static and dynamic characteristics between the damaged and undamaged models are compared. The results indicate that when the extent of damage increases, the change rate of drag coefficient at larger AoA (angle-of-attack) is greater than that at the smaller AoA. On the contrary, the change rate of lift coefficient at larger AoA is smaller than that at smaller AoA. Meanwhile, the absolute value of the static pitch moment decreases, but the absolute value of the roll moment increases. Damage causes a change in the absolute values of the pitch and roll dynamic derivatives, and the dynamic derivatives do not vary monotonically with the increase of AoA. The turning point occurs at about $\alpha$ = 5°. The areas of the hysteresis loops of the pitch-roll coupling moment increase, which makes the dynamic coupling characteristic between the pitch and roll directions increase. Finally, the maximum allowable damage extent of the missile wing that can achieve static trim is obtained and validated by controlling the deflection of the four rudders.

Theoretical research on dynamic modeling of rigid–flexible coupling system with double joint folded wing

For meeting the requirements of tactical missiles seeking miniaturized launch devices for storage, transportation, and launch, a tube-launched missile wing is adopted, which folds before launch and quickly unfolds after launch. As a structure installed on the missile body to generate the required aerodynamic force for manipulating the missile, the tube-launched missile wing can effectively stabilize the missile’s flight attitude. At present, most research on the unfolding mechanism of missile folding wings is focused on one-time folding. When the wingspan is large, multiple folding is required to meet the launch requirements of modern tube-launched missiles. Therefore, this article designs a dual-joint folding wing deployment mechanism and studies the rigid–flexible coupling dynamic modeling and related technologies of folding wings based on this structure. Based on the inertial coupling between large-scale rigid body motion and structural flexible deformation, the folding wing breaks through the element convergence of the model and achieves the applicability of the structural model through zero-order approximation model analysis and other technologies. Simulation results show that the hybrid coordinate method can fully and accurately display the vibration information of flexible folding wings. At different speeds, the first-order coupling model is more advanced than the zero-order coupling model. In addition, increasing rotational speed, increasing wing thickness, and reducing wing span length can effectively increase the fundamental frequency of wing flutter. The structural design of folding wings has shown important reference significance.

Research on the motion characteristics of the opening process of double joint folding wings

The folding wing mechanism is a system developed for the miniaturization of missile wings to facilitate box (cylinder) storage, transportation, and launching, saving the storage and transportation space of the missile and improving the combat capability of the weapon system. In order to solve the stability problem of the dynamic opening process of wings during missile flight, a theoretical model of a double joint folding wing is proposed. Based on dynamic theory, this paper studies the dynamic process of the folding wing system undergoing rotational opening motion under different initial forces. The influence of the initial angular velocity required for folding wing deployment on the deployment process is analyzed through structural model simulation. Finally, the effect of the folding wing opening action on the flight motion of the full missile is studied, and the dynamic opening process of the folding wing system is simulated by physical experiments. The research results show that the opening motion of the folding wing is sensitive to the initial conditions; during the deformation process, a change in motion type was observed, and the aeroelastic response at certain folding angles may exhibit different motion types. The main impact on the flight motion of the full missile is the change in the aerodynamic force of the missile.

Silver Lining Study: Missileer Fatigue Mitigation During 2020 Coronavirus Pandemic.

The U.S. Air Force's Intercontinental Ballistic Missile (ICBM) force stands ready to launch weapons 365 days per year. Since its inception, missileers vigilantly operate launch consoles on a 3-day cycle: minimum 24-hour alert-shift/24-hour travel-admin/24-hour off, leading to concerns that health, morale, and alertness are chronically impacted. In 2020, a Missileer Occupational Health Assessment (OHA) revealed 76% of respondents struggle with being rested for duty and 29% of respondents never feel adequately rested for duty. Later that year, 20th Air Force initiated long-duration operations to safeguard from the SARS CoV-2 (COVID-19) Pandemic, resulting in increased operations tempo, and exacerbating crew fatigue.341st Operations Group and 341st Medical Group at Malmstrom Air Force Base enacted interventions to mitigate crew fatigue and support continued readiness during pandemic operations. They recorded, analyzed, and compiled findings in this report, including recommendations for long-term ICBM operations at Missile Wings. All participants were Nuclear and Missile Operations Officers, or missileers, were continuously evaluated with qualitative and quantitative measures to ensure safety of the force during a period of unprecedented change. Interventions implemented and evaluated during the 9-month period included: environmental modifications, scheduling changes, and crew education on fatigue management, nutrition, anticipatory sleep preparation, and fitness. Most notably, the 341st Operations Group examined various 3-person and 4-person shift-length and alert duration schedules. Psychomotor vigilance testing results validated safety of operators and delta between pre- and post-shift measurements. Crew force readiness trends were analyzed for force-health awareness. Pre- and post-OHA results were compared for subjective changes. Fatigue and health-related outcomes were collected from a safety monitoring effort during standard and COVID-19 operations at 341st Missile Wing. Findings from qualitative and quantitative data indicate the optimal schedule is a 3-week cycle:7-day alert/7-day recovery/7-day training-administrative utilizing 4-member or 3-member crews for low tempo operations. Crews experimented with shift-lengths of 24hrs-on/24hrs-off, 16hrs-on/8hrs-off, and 12hrs-on/12hrs-off. Maximum safe alert duration is 7 days due to task fatigue onset between 8 and 10 days. Short and long duration Duties Not to Include Flight (DNIF) (also known as Duties Not to Include Alert (DNIA) among missileers) rates decreased from the first to last month of the period by 74.6% and 79.2%, respectively. The number of alerts missed per month decreased 86% from baseline. The 2021 OHA found a 7% decline in members seeking separation, and absence of sleep, fatigue, and physical or mental health as missileer concerns. This analysis has identified a sustainable alert rotation of 7/7/7 with emphasis on protected recovery and training time and has been continued after concluding pandemic operations, creating consistent schedule stability where there once was none. If executed properly, this alert rotation, regardless of shift-length selected, has potential to improve trust between crews and leadership, provides adequate recovery time between alerts to maintain health, and improves wellness, family stability, morale, unit cohesion, and crew force retention. Notably, all Air Force Global Strike Missile Operations Groups adjusted scheduling practices to align with these findings.

The effects of different wing configurations on missile aerodynamics

In the present study, missile aerodynamic analysis is performed using different wing config-urations at subsonic and transonic speeds. The wing is critical component in point of aero-dynamic efficiency for a missile that speed is especially closer to transonic level because of flow separation. Flow on the wings may adversely effect tailfins of missile at high speed since it may cause vortex generation and flow disturbances. There are few studies that investigate the missile wing using different configurations at critical speeds when examined the previ-ous studies. Therefore, in this study, three different wing configurations are investigated and aerodynamic performance is compared with each other at 0.7 and 0.9 Mach numbers and 5° angle of attack (AoA). In beginning of this study, missile model with only tailfins is selected from previous study that contains experimental data. Because the experimental data for the selected missile model are available at supersonic speeds, the aerodynamic analysis to verify the solutions is carried out at supersonic speeds. After wing is mounted to the selected missile, aerodynamic analysis is carried out using three different wing configurations that are Tapered Leading Edge, Tapered Trailing Edge, and Double Tapered wings. Lift to drag ratio (CL/CD) is calculated to compare wing configurations and it is concluded that Tapered Leading Edge wing configuration shows higher performance then other wing configurations. CL/CD values are 2.327, 2.306, 2.303 at 0.7 Mach number and 2.45, 2.429, 2.423 at 0.9 Mach number for Tapered Leading Edge, Tapered Trailing Edge, and Double Tapered, respectively. When the results are compared each other, CL/CD values at 0.9 Mach number is higher about % 5.28, %5.33 and %5.21 than the CL/CD values at 0.7 Mach number for missile with Tapered Leading Edge, Tapered Trailing Edge, and Double Tapered, respectively.

Ideal trajectory generation method of missile formation based on Gaussian pseudo-spectrum method

With the increasingly complex combat environment and the diversity of combat tasks, important military targets are surrounded by multi-layer defense. Modern information warfare pays more attention to system confrontation and multi-system cooperative operations,and missile formation to attack targets is the main research content of this paper. For missile wing ideal trajectory generation problems, ensure the attack Angle of time and meet the design requirements, Gaussian pseudo-spectral method is used to generate guarantee missile attacks within the specified time shot on target, at the same time meet the attack Angle constraint at the end of the ideal trajectory, the more missiles cooperative engagement problems into the ideal trajectory tracking problem, avoid the estimate of the remaining time of flight.The main contents of this paper are as follows:(1) Gaussian pseudo-spectral method is used to generate the ideal trajectory model.(2)Use Matlab to simulate the established model.

Optimization Design of the Wing of Cruise Missile on the Basis of the Vortex Lattice Approach

Aerodynamic performance of missile wings is a vital factor affecting the endurance and combat effectiveness of cruise missiles. Given the current situation that the aerodynamic design of conventional cruise missile wings is less and there is more room for improvement, a missile wing shape with a winglet is considered designed to improve the cruise characteristics of cruise missiles effectively. Firstly, the EPPLER 66 is selected as the wing shape by the combination of vortex surface element approach batch calculation and CFD simulation analysis. According to the dimensional parameters of a typical patrol wing, the design variables and corresponding optimization intervals are determined, and the aerodynamic optimization is carried out by combining the genetic approach as well as the vortex lattice approach. CFD is used to verify the aerodynamic performance of the optimized wing. The calculation outcomes reveal that the optimized wing has a ratio of lift-to-drag of 23.43, which is 12.32% higher than that of the conventional flat wing.

Time-Varying Aeroelastic Modeling and Analysis of a Rapidly Morphing Wing

Advanced rotational variable-swept missile wings require the ability to rapidly deploy, retract and reach the designated position. Therefore, the establishment of an effective time-varying aeroelastic model of a rotating missile wing is the prerequisite for performing transient response analysis during the rapid morphing process. In this paper, the finite element model of the wing at the fixed configuration is combined with the floating frame method to describe the small elastic deformations and large rigid-body displacements of the wing, respectively. Combining the structural dynamic model with the supersonic piston theory, a nonlinear and time-varying aeroelastic model of a missile wing undergoing the rapid morphing process is established. A method for the real-time determination of the time-varying lifting surface during morphing is discussed. Based on the proposed aeroelastic equations of motion, the flutter characteristics of the wing at different sweep angles are obtained. The influences of the actuator spring constant, the damping ratio during the morphing and the post-lock stages, as well as the velocity quadratic term in the aeroelastic equations, on the transient responses of the system are studied. The simulation results show that the flutter characteristics of the wing are greatly influenced by the sweep angle. Moreover, the jumping phenomenon in flutter speed due to the switching of flutter modes is found with the increase of the sweep angle. The morphing simulations demonstrate that the transient aeroelastic responses mainly occur in the post-lock stage, so much more attention needs to be focused on the post-lock vibrations. In addition, under the given simulation parameters, the nonlinear quadratic velocity term has little effect on the transient responses of the system. This study provides an efficient method for predicting the transient aeroelastic responses of a rotational variable swept wing.

Design, molding, manufacturing and testing of wave-transparent functional composite missile wings

In this paper, a fiber-reinforced resin matrix composite missile wing with a wave-transparent function is studied, which has the function of wave-transparent and meets the requirements of the mechanical properties of the missile during flight, and the missile wing structure is made of aluminum alloy, carbon fiber composite material, and glass fiber composite material, and the weight reduction is about 30.3% compared with the overall aluminum alloy structure of the missile wing. In the design process, the finite element simulation method is used, the plastic deformation of aluminum alloy is fully considered, and the antenna is built into the airfoil of glass fiber composite material, which successfully realizes the wave-transparent function of the missile wing and provides a new design idea for the composite wing.

Design and Analysis of a Deployment Mechanism with Clearance Compensation for High Stiffness Missile Wings

The deployment performance of the unfolded wing determines whether the winged missiles can fly normally after being launched, infecting the attack performance of the winged missiles. The paper proposes a new deployment mechanism with clearance eliminator. Based on the slider-crank principle, the proposed deployment mechanism achieves fast and low-impact deployment of the wings. The proposed clearance eliminator with shape memory alloy (SMA) effectively eliminates the clearance of the sliding pair and improves the support stiffness and stability of the deployed wing. The collision characteristics and the clearance elimination are studied for the deployment mechanism. The influence of the collision force on the motion state of the wing during the deployment is analyzed. The static stiffness of the wing under the clearance state and the deformation is analyzed. The dynamic stiffness under the catapult clearance elimination state is modeled based on the fractal geometry and contact stress theory. The relationship between the locking force and the support stiffness is revealed. The kinetic simulation is used to analyze the motion response during the action of the deployment mechanism. Modal analysis, harmonic response analysis, and random vibration analysis were conducted for the whole wings. A prototype was developed to verify the ejection performance of the wing according to the input load characteristics. The dynamic stiffness of the unfolded wings is tested by the fundamental frequency experiments to verify the performance of the clearance elimination assembly. The experimental results show that the designed deployment mechanism with clearance compensation achieves fast ejection and high stiffness retention of the missile wing.

ВИЗНАЧЕННЯ ПОТРІБНОГО ПИТОМОГО НАВАНТА-ЖЕННЯ НА КРИЛО РАКЕТИ

Wing loading severely affects the mass of a missile as well as its flight performance. For airplanes this parameter must not exceed allowable values calculated from different requirements especially related to such cases as lift-off, landing, cruise flight and aircraft maneuverability. For missiles wing loading is determined considering launch conditions and providing the necessary maneuverability. Appropriate estimation of wing loading at the initial design stages guarantees the minimal mass of an aircraft with all tactical requirements met.Review of available literature, related to missile design, has shown that the problem of optimal wing loading estimation contains lengthy and quite approximate analytical expressions.This article is dedicated to the development of a missile wing loading estimation technique that provides minimal propellant mass fraction and total mass of an aircraft while meeting tactical requirements.Impact of wing loading onto propellant mass fraction, maximal maneuverability and total mass of a missile is considered. The algorithm of optimal wing loading estimation, which provides necessary tactical characteristics of a missile being designed, is proposed. We define simple polynomial approximations of both the trajectory and the velocity profile. Further analysis is being conducted using two considerations: for an air-based missile the value of wing loading has to provide flight during launch without fall movement as well as the maximal maneuverability at the moment when a missile intercepts the target.It is shown that for the wing loading in the range from 300 to 1000 kg/m2 the propellant mass fraction changes rapidly, and inaccurate selection of wing loading may lead to obtaining of an incorrect value of propellant mass fraction. For maximal maneuverability less than 40, inappropriate selection of wing loading may cause significant numerical error. Analysis of relation between wing loading and total mass of a missile revealed that there is a critical value of wing loading which depends on initial data and represents the low limit of an acceptable range.

Visual dynamics simulation for adapters separation of missile launching

The process of the adapter separating from the missile plays an important role in the safety and reliability of missile launching, which involves complex dynamics calculations. In order to guarantee the safety of separating the adapter from the missile and successfully launching the missile, this article was concerned with the whole process of the adapter separating from the missile launching system and presented a novel calculation method and visual simulation software to analyze the dynamics process of adapter separation. The dynamics models including the coordinate system, dynamics and kinematic equations, and the forces acting on the adapter are established. The random factors and their distribution law that greatly influence the separation process are analyzed, the confidence intervals of the separation reliability are estimated by using the Monte Carlo method, and an algorithm for calculating the distance between the adapter, missile wing, and tail rudder is proposed. Visual simulation software for analyzing the separation process is designed, which gathers modeling, solving, and analytical processing. Finally, a numerical simulation result is given to verify the validity and correctness of the proposed method and software and analyze the factors affecting the separation reliability of the separating process of adapters for missile launching. The method and software presented in this paper can provide support for the safety design of missile adapter separation.

Analysis of pressure characteristics under laminar and turbulent flow states inside the pilot stage of a deflection flapper servo-valve: Mathematical modeling with CFD study and experimental validation

Electro-hydraulic servo-valves are widely used components in the mechanical industry, aerospace and aerodynamic devices which precisely control the airplane or missile wings. Due to the small size and complex structure in the pilot stage of deflection flapper servo-valves, accurate mathematical models for the flow and pressure characteristics have always been very difficult to be built. In this paper, mathematical models for the pilot stage of deflection flapper servo-valve are investigated to overcome some gaps between the theoretical formulation and overall performance of the valve by considering different flow states. Here, a mathematical model of the velocity distribution at the flapper groove exit is established by using Schlichting velocity equations for in-compressible laminar fluid flow. Moreover, when the flow becomes turbulent, a mathematical model of pressure characteristics in the receiving ports is built on the basis of the assumption of the collision between the liquid and the jet as the impact of the jet on a moving block of fluid particles. To verify the analytical models for both laminar and turbulent flows, the pressure characteristics of the deflection flapper pilot stage are calculated and tested by using numerical simulation and experiment. Experimental verification of the theory is also presented. The computed numerical and analytical results show a good agreement with experimental data.

Study of linear ablative rate of D6AC steel wing used on supersonic missile

The D6AC steel wing used on supersonic missile is the object in this study. Its service environment was generated and simulated. The ablation experiment of D6AC steel missile wing was carried out under different parameters of flow field. The ablation process of D6AC steel wing was studied and analyzed under the supersonic aerodynamic heating environment. The results show that the ablation process of D6AC steel missile wing could be divided into three stages: aerodynamic heating, oxidation reaction, and shear stripping. The influence factors of the D6AC, steel wing ablation include the total temperature, Mach number, oxygen content, and water content. The higher the total temperature is, the more early the initial ablation time of wing is. The linear ablative rate of the D6AC steel wing is the result of the interaction of the Mach number, oxygen content, and water content. The higher the Mach number is, the larger the oxygen content is, and the larger the linear ablative rate of missile wing is. The influence of water content on the linear ablative rate of D6AC steel wing is the opposite.

Assessment of the performance of a cruise missile wing depending on design parameters at various flight conditions

The design of a cruise missile wing is a multidisciplinary work, since the wing has to provide enough aerodynamic forces while being structurally safe. In order to meet the design requirements and achieve a successful design, several parameters of the wing need to be investigated at various flight conditions. In this paper, lift forces and flutter speeds are obtained for different wing geometries, at various altitudes and speeds. The design of experiments is constructed by changing the several variables of the wing to have better understanding of the effect of design parameters. The aerodynamic analyses for predicting the lift force are performed by using FLUENT and the aeroelastic analyses for the determination of the flutter speed are conducted by using ZAERO package programs.

Realizacija servosistema za vođenje rakete zasnovanog na primeni mikrokontrolera i beskontaktnog senzora položaja krila rakete

This paper describes the procedure for the realization of servo systems for the guided missiles. The realized servo system consists of four servo units for driving the missile wings. The servo units are independent of the each other and they are controlled by a 32-bit microcontroller STM32F407. Contactless high resolution magnetic encoders are used as the position sensors of the missile wings. The communication between the microcontroller and the magnetic encoders is realized via the serial port using the SPI protocol. DC motors are used as actuators. The actuator's energy converters are realized in a Hbridge configuration. These converters are controlled by the corresponding pulse width modulation (PWM) signal from the microcontroller. This servo system with four independents servo units makes possible to rotate the missile wings according to the guidance algorithm and at the same time according to the algorithm for the missile rolling control. The steering structures of the servo units are based on the application of the PI controllers.

Application of Free Interface Modal Synthesis Method on Modal Test Boundary Transformation of One Missile Wing

The large deviation between test results and real conditions caused by the coupling of test fixture is unavoidable. The research on free interface modal synthesis method can gain the modal test boundary transformation of one missile wing. For this purpose, the low order normal mode spectrum matrix and mode shapes was gained by free modal test, the extra high orders modes can be obtained through FEM analysis, while the boundary constrains was equivalent as a kind of sub-structure, and the modal synthesis was used to extract the constrained modal parameters. The experiment results are consistent with numeric calculation results very well, which indicated that this extracting method is feasible.

Milling research and tool selection design of SiC14Cu4Mg0.5Si based on Aluminium matrix 2A14

The selection of milling tools for SiC14Cu4Mg0.5Si based on Aluminium matrix 2A14 was analyzed, and the factors that affect the efficiency of the milling were discussed. The SiC14Cu4Mg0.5Si was designed for use on the moon landing vehicle or missile wings, but the hardness of aluminium-silicon carbide composite material was very high, much higher than the general hardness of cemented carbide, which will bring many difficulties in the aluminium-silicon carbide composite material processing. The chemical compositions of SiC14Cu4Mg0.5Si were analyzed. A new selected indexable cutter was designed to mill SiC14Cu4Mg0.5Si. The structure design of milling cutter was different from the conventional milling cutter, breaking the previous limitations to a certain extent, pioneering the idea. The tool material wear was detected by experiments. The mechanical and physical properties of SiC14Cu4Mg0.5Si were also tested. SiC14Cu4Mg0.5Si exhibited different surface quality characteristics under different milling tools.

Nonlinear flutter and active control of the supersonic two-dimensional missile wings with a freeplay

The nonlinear flutter characters and active control of the supersonic two-dimensional missile wings with a freeplay are investigated. Adopting the third-order piston, the nonlinear aeroelastic dynamic equations with freeplay nonlinearity are derived, nonlinear flutter character analysis of the systems are obtained by hopf bifurcation theory, harmonic balance method and numerical method respectively. The design method based on differential geometry and quadratic optimal control for nonlinear system controller is used to postpone the critical flutter speed. The sliding mode variable structure control for nonlinear system controller is used to stabilize the nonlinear flutter. Then, the influence of the controller and the parameters are analyzed. The simulation result shows that the two control methods are effective ways in control of nonlinear flutter. At last, the dynamical response with the base excitation disturbance are calculated, the single-periodic motion, the multi-periodic motion, the quasi-periodic motion and the chaotic motion are obtained.