Artykuły w czasopismach: „Launch Vehicle Model”

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You, Ming, Qun Zong, Bailing Tian i Fanlin Zeng. "Nonsingular terminal sliding mode control for reusable launch vehicle with atmospheric disturbances". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, nr 11 (8.05.2017): 2019–33. http://dx.doi.org/10.1177/0954410017708211.

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The controller design for reusable launch vehicles is challenging due to enormous amounts of model parameter uncertainties and atmospheric disturbances. This paper first derives six-degree-of-freedom model of a reusable launch vehicle with atmospheric disturbances. Next, four kinds of atmospheric disturbances are introduced and wind models are established respectively. For attitude control of the reusable launch vehicle, a nonsingular terminal sliding mode controller is designed with stability guaranteed. Finally, simulation results show a satisfactory performance for the attitude tracking of the reusable launch vehicle with atmospheric disturbances.

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Gibson, Denton, Waldemar Karwowski, Timothy Kotnour, Luis Rabelo i David Kern. "The Relationships between Organizational Factors and Systems Engineering Process Performance in Launching Space Vehicles". Applied Sciences 12, nr 22 (14.11.2022): 11541. http://dx.doi.org/10.3390/app122211541.

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The launch vehicle industry has long been considered a pioneering industry in systems engineering. Launch vehicles are large complex systems that require a methodical multi-disciplinary approach to design, build, and launch. Launch vehicles are used to deliver payloads—such as humans, robotic science missions, or national security payloads—to desired locations in space. Previous research has identified deficient or underperforming systems engineering as a leading contributor to launch vehicle failures. Launch vehicle failures can negatively affect national security, the economy, science, and society, thus highlighting the importance of understanding the factors that influence systems engineering in launch vehicle organizations in the United States. The purpose of this study was to identify and evaluate the relationships between organizational factors and systems engineering process performance. Structural equation modeling was used to develop a model of the relationships of these factors and test hypotheses. The results showed that organizational commitment, top management support, the perceived value of systems engineering, and systems engineering support significantly influence systems engineering process performance in the launch vehicle industry. Implications of this study for improving the performance of systems engineering in launch vehicle organizations are discussed.

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Hladkyi, Ye H., i V. I. Perlyk. "How Yuzhnoye develops models for flight safety index evaluation for the case of a rocket failure during the flight". Kosmičeskaâ tehnika. Raketnoe vooruženie 2023, nr 1 (12.05.2023): 14–30. http://dx.doi.org/10.33136/stma2023.01.014.

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Safety of the up-to-date rocket and space complexes remains a topical problem for the developers of rocket and space technology. The integral component of this problem along with the safety of operations during launch vehicle ground pre-launch processing is organization of flight safety. The basic task of this rocket and space complexes safety component is to prevent or minimize serious consequences in case of launch vehicle failure in the flight leg, after all such accidents can cause damage to the population and facilities (including personnel and facilities of the ground complex), located along the flight paths. It is shown that the flight safety assurance of the launch vehicle is based on the experience of combat missile systems. Flight safety during the launch vehicle launches is provided by laying flight paths through sparsely populated (unpopulated) territories and using special onboard flight safety systems. This system limits the size of impact zones of emergency launch vehicle and its debris by emergency engine shutdown. Recently flight safety process is organized based on the acceptable risk concept. It is based on a risk assessment for the ground-based facilities and people, and it should not exceed the established standards. Such approach requires development and upgrading of the mathematical models of risk assessment in case of launch vehicle failure in the flight phase. Formation of the risk-oriented approach to flight safety in Yuzhnoye SDO is shown. Key moment in this process is to develop the separate structural unit, which started working on rocket and space complexes flight safety assurance and analysis. The basic model for assessing the risks of damage to facilities and people is analyzed, using the maximum impact zone of an emergency launch vehicle, which is realized in case of loss of control and flight safety system activation. The main directions of the basic model improvement are shown, which led to the development of a number of new original models of flight safety assessment in the Yuzhnoye SDO. First of all, the developed models take into account the flight safety system specifics, which are used to equip the launch vehicles, developed by Yuzhnoye SDO: criteria of activation, blocking of the engine emergency shutdown in the initial flight phase and Fe functional. Such models allow to take into account the different nature of emergency situations in the launch vehicle flight phase and ways of their representation, representation of the damage areas of facilities in the form of convex polygons, possible fragmentation of the emergency launch vehicle at the free- fall leg etc. The developed models have found wide application in the practice of assessing flight safety indicators in the Yuzhnoye SDO projects. Key words: launch vehicle; acceptable risk; launch vehicle failure in the flight phase; flight safety system; emergency launch vehicle impact zone; risk of damage to facilities; collection risk.

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Pu, Pengyu, i Yi Jiang. "Assessing Turbulence Models on the Simulation of Launch Vehicle Base Heating". International Journal of Aerospace Engineering 2019 (22.08.2019): 1–14. http://dx.doi.org/10.1155/2019/4240980.

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Launch vehicles suffer from severe base heating during ascents. To predict launch vehicle base heat flux, the computational fluid dynamics (CFD) tools are widely used. The selection of the turbulence model determines the numerical simulation results of launch vehicle base heating, which may instruct the thermal protection design for the launch vehicle base. To assess performances, several Reynolds-averaged turbulence models have been investigated for the base heating simulation based on a four-nozzle launch vehicle model. The finite-rate chemistry model was used for afterburning. The results showed that all the turbulence models have provided nearly identical mean flow properties at the nozzle exit. Menter’s baseline (BSL) and shear stress transport (SST) models have estimated the highest collision pressure and have best predicted base heat flux compared to the experiment. The Spalart-Allmaras (SA) model and the renormalization group (RNG) model have performed best in temperature estimation, respectively, in around r/rb=0~0.2 and r/rb=0.6~1. The realizable k‐ε (RKE) model has underestimated the reverse flow and failed to correctly reflect the recirculation in the base region, thus poorly predicted base heating. Among all the investigated turbulence models, the BSL and SST models are more suitable for launch vehicle base heating simulation.

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da Cás, Pedro L. K., Carlos A. G. Veras, Olexiy Shynkarenko i Rodrigo Leonardi. "A Brazilian Space Launch System for the Small Satellite Market". Aerospace 6, nr 11 (12.11.2019): 123. http://dx.doi.org/10.3390/aerospace6110123.

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At present, most small satellites are delivered as hosted payloads on large launch vehicles. Considering the current technological development, constellations of small satellites can provide a broad range of services operating at designated orbits. To achieve that, small satellite customers are seeking cost-effective launch services for dedicated missions. This paper deals with performance and cost assessments of a set of launch vehicle concepts based on a solid propellant rocket engine (S-50) under development by the Institute of Aeronautics and Space (Brazil) with support from the Brazilian Space Agency. Cost estimation analysis, based on the TRANSCOST model, was carried out taking into account the costs of launch system development, vehicle fabrication, direct and indirect operation cost. A cost-competitive expendable launch system was identified by using three S-50 solid rocket motors for the first stage, one S-50 engine for the second stage and a flight-proven cluster of pressure-fed liquid engines for the third stage. This launch system, operating from the Alcantara Launch Center, located at 2 ∘ 20’ S, would deliver satellites from the 500 kg class in typical polar missions with a specific transportation cost of about US$39,000 per kilogram of payload at a rate of 12 launches per year, in dedicated missions. At a low inclined orbit, vehicle payload capacity increased, decreasing the specific transportation cost to about 32,000 US$/kg. Cost analysis also showed that vehicle development effort would claim 781 work year, or less than 80 million dollars. Vehicle fabrication accounted for 174 work year representing less than 23 million dollars per unit. The launch system based on the best concept would, therefore, deploy small satellite constellations in cost-effective dedicated launches, 224 work year per flight, from the Alcantara Launch Center in Brazil.

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Adelfang, S. I., O. E. Smith i G. W. Batts. "Ascent wind model for launch vehicle design". Journal of Spacecraft and Rockets 31, nr 3 (maj 1994): 502–8. http://dx.doi.org/10.2514/3.26467.

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Wang, J. T., G. Y. Hang, H. M. Shen, Z. Y. Liu, H. J. Xue, T. Wang i W. Yu. "Numerical Simulation of Shock Wave Damage to Medium-Range and Long-Range Targets". Journal of Physics: Conference Series 2478, nr 2 (1.06.2023): 022002. http://dx.doi.org/10.1088/1742-6596/2478/2/022002.

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Abstract In order to accurately analyze the effect of a shock wave on a missile launch vehicle as a whole and on its components when the launch vehicle is at a medium or long distance from the detonation center, a simulation method based on the empirical algorithm and numerical analysis was carried out in this study. The method significantly reduced the computational cost while ensuring computational accuracy. Based on the simulation method, a finite element model for a typical missile launch vehicle was established that consisted of 2.5 million elements. Based on the structured arbitrary Lagrangian-Eulerian method and the fluid-structure coupling algorithm, it took the model only a few hours to simulate the second-level physical process. Next, a shock wave load model was built with a 2000 kg TNT equivalent detonation condition, and the degree of damage to the launch vehicle within 25–45 m from the detonation center was analyzed. The results showed that the joint action of the shock wave overpressure and the dynamic pressure was the main source of damage. Specifically, the missile launcher, the cab, and the fuel tank were the key vulnerable parts. The effective damage radius of the 2000 kg TNT equivalent detonation to the missile launch vehicle was 35 m.

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Golubek, A. V., i N. M. Dron'. "Launch Vehicle Rendezvous to Catalogued Orbital Debris while Injecting into Highly-Inclined Orbits". Nauka ta innovacii 16, nr 6 (12.06.2020): 46–55. http://dx.doi.org/10.15407/scin16.06.046.

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Introduction. A constant increase in the amount of space debris already constitutes a significant threat to satellites in nearEarth orbits, starting with the trajectory of their launch vehicle injection. Problem Statement. Design and development of various modern methods of protection against space debris requires knowledge of the statistical characteristics of the distribution of the kinematic parameters of the simultaneous motion of a launch vehicle injecting satellite and a group of space debris objects in the area of its trajectory. Purpose. Development of a mathematical model of a launch vehicle rendezvous with a group of observable orbital debris while injecting a satellite into near-earth orbits with an altitude of up to 2100 km and an inclination from 45 to 90 degrees. Materials and Methods. The following methods are used in the research: analysis, synthesis, comparison, simulation modeling, statistical processing of experimental results, approximation, correlation analysis, and the least squares method. Results. The simultaneous motion of a launch vehicle and a group of space debris objects has been studied. The distributions of relative distance, relative velocity, angle of encounter, and moments of time of approach of a launch vehicle to a group of the observed space debris at a relative distance of less than 5 km have been obtained. The dependence of the average rendezvous concentration on the distribution of space debris across the average altitude of the orbit and the inclination of the target orbit of the launch vehicle has been determined. The dependence of the average probability of rendezvous in the launch on the inclination of the target orbit, the number of orbital debris, and the relative distance of the rendezvous has been determined. Conclusions. The obtained mathematical model of rendezvous of a launch vehicle with a group of observed orbital debris can be used while designing means of cleaning the near-Earth space and systems to protect modern satellite launch vehicles from orbital debris. In addition, the results of the research can be used to assess the impact of unobserved orbital debris on the flight of a launch vehicle.

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Golubek, A. V., i N. M. Dron'. "Launch Vehicle Rendezvous to Catalogued Orbital Debris while Injecting into Highly-Inclined Orbits". Science and innovation 16, nr 6 (listopad 2020): 46–55. http://dx.doi.org/10.15407/scine16.06.046.

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Introduction. A constant increase in the amount of space debris already constitutes a significant threat to satellites in nearEarth orbits, starting with the trajectory of their launch vehicle injection. Problem Statement. Design and development of various modern methods of protection against space debris requires knowledge of the statistical characteristics of the distribution of the kinematic parameters of the simultaneous motion of a launch vehicle injecting satellite and a group of space debris objects in the area of its trajectory. Purpose. Development of a mathematical model of a launch vehicle rendezvous with a group of observable orbital debris while injecting a satellite into near-earth orbits with an altitude of up to 2100 km and an inclination from 45 to 90 degrees. Materials and Methods. The following methods are used in the research: analysis, synthesis, comparison, simulation modeling, statistical processing of experimental results, approximation, correlation analysis, and the least squares method. Results. The simultaneous motion of a launch vehicle and a group of space debris objects has been studied. The distributions of relative distance, relative velocity, angle of encounter, and moments of time of approach of a launch vehicle to a group of the observed space debris at a relative distance of less than 5 km have been obtained. The dependence of the average rendezvous concentration on the distribution of space debris across the average altitude of the orbit and the inclination of the target orbit of the launch vehicle has been determined. The dependence of the average probability of rendezvous in the launch on the inclination of the target orbit, the number of orbital debris, and the relative distance of the rendezvous has been determined. Conclusions. The obtained mathematical model of rendezvous of a launch vehicle with a group of observed orbital debris can be used while designing means of cleaning the near-Earth space and systems to protect modern satellite launch vehicles from orbital debris. In addition, the results of the research can be used to assess the impact of unobserved orbital debris on the flight of a launch vehicle.

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Peng, Bo, Cheng Ma, Guodong Wang, Fengyan Hu, Ke Mei i Jian Yang. "An aerodynamic surrogate model of launch vehicle based on relevance vector machine". Journal of Physics: Conference Series 2181, nr 1 (1.01.2022): 012021. http://dx.doi.org/10.1088/1742-6596/2181/1/012021.

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Abstract In the process of launch vehicle multidisciplinary design optimization, aerodynamic calculation takes a long time, which affects the overall design cycle. In order to solve the above problems, based on the idea of machine learning, this paper constructs the surrogate model of relevance vector machine and calculates the aerodynamic coefficients of launch vehicles quickly. Firstly, the aerodynamic model of launch vehicle is established, and the orthogonal design method is used to generate test sample points. Then, the aerodynamic coefficients of the sample points are calculated by using Fluent software, and the training data of the surrogate model are obtained. On this basis, the relevance vector machine model is trained with training data, generating correlation vector machine agent model. Finally, the calculation accuracy of the surrogate model is evaluated by simulation, and the feasibility and validity of the method are verified.

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Song, Haryong, i Yongtae Choi. "Distributed multiple model extended information filter with unbiased mixing for satellite launch vehicle tracking". International Journal of Distributed Sensor Networks 14, nr 4 (kwiecień 2018): 155014771876926. http://dx.doi.org/10.1177/1550147718769263.

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A distributed extended information filter-based interacting multiple model estimator with unbiased mixing is proposed for satellite launch vehicle tracking. In this problem, multiple heterogeneous sensors such as radars, telemetry systems receiving onboard Global Positioning System—inertial navigation system data, and electro-optical targeting systems are used. The extended information filter is used for nonlinear estimation dealing with ballistic model and spherical coordinate observation. The multiple Markov switching models comprise thrusting and coasting modes having different state vector dimensions for the launch vehicle. To effectively combine both state vectors, an unbiased mixing technique is applied and then the distributed extended information filter integrates local states and information matrix contributions. Hence, the proposed algorithm takes into account both heterogeneity of tracking sensors and multiplicity of vehicle’s dynamic model. We prove the superiority of the proposed algorithm by conducting Monte Carlo simulation with nominal trajectory data of Korea Space Launch Vehicle-1. Comparative simulation results demonstrate that the performance of the proposed method has been improved in vehicle’s position root mean square error.

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Liu, Bing, Xiaohan Chen, Enyi Li i Guigao Le. "Numerical Analysis on Water-Exit Process of Submersible Aerial Vehicle under Different Launch Conditions". Journal of Marine Science and Engineering 11, nr 4 (15.04.2023): 839. http://dx.doi.org/10.3390/jmse11040839.

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To study the influence of launch conditions and wave interference on the stability of submersible aerial vehicles at the water–air interface, a coupling model for water-exit motion of submersible aerial vehicles was established by using the RNG k-ε turbulence model and VOF method. The water-exit processes of submersible aerial vehicles under different initial inclination angles and velocities were numerically simulated and the effects of initial inclination angle and velocity on the water-exit motion of submersible aerial vehicles were obtained. Based on the response surface function theory, a mathematical model for the motion stability of submersible aerial vehicles at the water–air interface was established, so that the submersible aerial vehicle’s pitch angle and velocity at the end of vehicle’s water-exit process, corresponding to any initial inclination angle and velocity, can be solved. The deviation between the simulated calculation result and the established fitting function model result was 2.7%. The minimum water-exit velocity of submarine aerial vehicles should be greater than 10.8 m/s. The research provides technical support for the trans-media motion stability analysis and hydrodynamic performance design of the submersible aerial vehicle.

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Han, I., i R. M. Brach. "Impact throw model for vehicle-pedestrian collision reconstruction". Proceedings of the Institution of Mechanical Engineers, Part D: Journal of Automobile Engineering 216, nr 6 (1.06.2002): 443–53. http://dx.doi.org/10.1243/09544070260137381.

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A planar model for the mechanics of a vehicle-pedestrian collision is presented, analysed and compared with experimental data. It takes into account the significant physical parameters of wrap and forward projection collisions and is suitable for solution using mathematics software or spreadsheets. Parameters related to the pedestrian and taken into account include horizontal distance travelled between primary and secondary impacts with the vehicle, launch angle, centre-of-gravity height at launch, the relative forward speed of the pedestrian to the car at launch, distance from launch to a ground impact, distance from ground impact to rest and pedestrian-ground drag factor. Vehicle and roadway parameters include post-impact, constant velocity vehicle travel distance, continued vehicle travel distance to rest with uniform deceleration and relative distance between rest positions of vehicle and pedestrian. The model is presented in two forms. The first relates the throw distance to the initial vehicle speed. The second, intended for reconstruction, relates the vehicle speed to the pedestrian throw distance. The first form is used as means of comparison of the model with selected sets of experimental data taken from the current literature, including a variational study using Monte Carlo simulation. The second (reconstruction) form is derived analytically not empirically and the parameters have physical interpretations. In order to obtain parameter values, direct calculation or the method of least squares can be used. A comparison of the reconstruction model with results of other reconstruction models is presented.

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Liu, Hai Jun, Xing Zhi Peng i Zhen Zhu Zou. "Effects of the Launch Speed on Hydrodynamic Force of the Underwater Vehicle Vertical Launch with the Gas Curtain". Advanced Materials Research 625 (grudzień 2012): 84–87. http://dx.doi.org/10.4028/www.scientific.net/amr.625.84.

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The vertical launch of the gas curtain is a new underwater launch technology. The gravity effect of the launch speed on hydrodynamic characteristics of the underwater vehicle vertical launching by using the gas curtain has been studied by adopting the multiphase VOF model and the standardturbulence model. The relationship between the launch speed and the shape of the underwater vehicle has been achieved by using the numerical simulation. The relationships between the launch speed and the hydrodynamic characteristics of the underwater vehicle vertically launching from the tube, navigating in water and exiting water have been investigated by using numerical simulation. The hydrodynamic characteristics of underwater vehicle vertical launching by using the gas curtain method are small. The effects of the launch speed on the hydrodynamic characteristics of the underwater vehicle vertical launching in the gas curtain are small.

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Li, Jiamin, Jian Zheng i Shibo Jing. "Influence of cavitation state and launch angle on water-exit process of vehicle based on moving domain method". Journal of Physics: Conference Series 2472, nr 1 (1.05.2023): 012028. http://dx.doi.org/10.1088/1742-6596/2472/1/012028.

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Abstract This paper, aiming at the water-exit process of the vehicle with cavitation, based on the moving domain method coupled with the SDOF (six degrees of freedom) solver, using VOF (volume of fluid) multiphase model and the Schnerr-Sauer cavitation model, carrying out numerical simulation research on multi-conditions. Comparing vertical launch vehicles with different water-exit Ca (cavitation number) cases, the sudden decrease in cavitation number is closely related to the cavitation state, and there is a maximum influent cavitation state of the cavitation collapse on the process of the water-exit. Comparing different initial launch angle cases, it is found that there is a most unstable launch angle, which makes the angular deflection of the vehicle the most severe during the water-exit process.

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Boglis, Ioana-Carmen, i Adrian M. Stoica. "Attenuation of the effects produced by the bending modes of a flexible launch vehicle using second order filters". Technium: Romanian Journal of Applied Sciences and Technology 2, nr 1 (7.01.2020): 48–55. http://dx.doi.org/10.47577/technium.v2i1.41.

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This article describes the application of an adaptive control algorithm for the atmospheric phase control of a launch vehicle. A dynamic model representing the pitch motion of the launch vehicle is introduced. As the bending modes can cause instability, the first threebending modes are modelled. A case study on design of the bending filters for a flexible launch vehicle using a model reference adaptive control is conducted to demonstrate its effectiveness.

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陈, 宜成. "Model Predictive Control Based Launch Vehicle Trajectory Optimization Method". Journal of Aerospace Science and Technology 08, nr 03 (2020): 49–59. http://dx.doi.org/10.12677/jast.2020.83007.

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Khoshnood, A., J. Roshanian i A. Khaki-Sedig. "Model reference adaptive control for a flexible launch vehicle". Proceedings of the Institution of Mechanical Engineers, Part I: Journal of Systems and Control Engineering 222, nr 1 (luty 2008): 49–55. http://dx.doi.org/10.1243/09596518jsce469.

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Sim, Chang-Hoon, Geun-Sang Kim, Dong-Goen Kim, In-Gul Kim, Soon-Hong Park i Jae-Sang Park. "Experimental and Computational Modal Analyses for Launch Vehicle Models considering Liquid Propellant and Flange Joints". International Journal of Aerospace Engineering 2018 (2018): 1–12. http://dx.doi.org/10.1155/2018/4865010.

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In this research, modal tests and analyses are performed for a simplified and scaled first-stage model of a space launch vehicle using liquid propellant. This study aims to establish finite element modeling techniques for computational modal analyses by considering the liquid propellant and flange joints of launch vehicles. The modal tests measure the natural frequencies and mode shapes in the first and second lateral bending modes. As the liquid filling ratio increases, the measured frequencies decrease. In addition, as the number of flange joints increases, the measured natural frequencies increase. Computational modal analyses using the finite element method are conducted. The liquid is modeled by the virtual mass method, and the flange joints are modeled using one-dimensional spring elements along with the node-to-node connection. Comparison of the modal test results and predicted natural frequencies shows good or moderate agreement. The correlation between the modal tests and analyses establishes finite element modeling techniques for modeling the liquid propellant and flange joints of space launch vehicles.

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Homsup, Pinyochon, Narathee, Weng i Homsup. "Dynamic Simulation of a UAV Moving on Launcher". Proceedings 39, nr 1 (21.01.2020): 26. http://dx.doi.org/10.3390/proceedings2019039026.

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This paper presents a dynamic simulation of Unmanned Aerial Vehicle (UAV) moving on a launcher. The UAV launcher consists of a metal rail that is positioned at a launch angle, a spring and a winch to be able to stretch the spring. Prior to launch the spring is stretched into necessary tension and secured with safety pins, then released to launch the UAV. The influence of aerodynamic lift and drag forces on the mathematical model of a UAV will be considered in detail. Mathematical model a UAV moving on a launcher represented by differential equations are solved using a MATLAB software. The results of calculations are represented by a graph which shows distance, velocity and acceleration of a UAV while moving on a launcher as a function of time.

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Kim, J. B., J. S. Sim, S. G. Lee, S. J. Shin, J. H. Park i Y. Kim. "Integrated one-dimensional dynamic analysis methodology for space launch vehicles reflecting liquid components". Aeronautical Journal 121, nr 1243 (11.07.2017): 1217–38. http://dx.doi.org/10.1017/aer.2017.56.

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ABSTRACTIn this paper, structural modelling and dynamic analysis methods reflecting the characteristics of a liquid propellant were developed for a pogo analysis. The pogo phenomenon results from the complex interaction between the vehicle structural vibration in the longitudinal direction and the propulsion system. Thus, for an accurate vibration analysis of a liquid propellant launch vehicle, both the consumption of the liquid propellant and the change in the stiffness reflecting the nonlinear hydroelastic effect were simultaneously considered. A complete vehicle structure, including the liquid propellant tanks, was analytically modelled while focusing on pogo. In addition, a feasible liquid propellant tank modelling method was established to obtain an one-dimensional complete vehicle model. With these methods, comparative studies of the hydroelastic effect were conducted. Evaluations of the dynamic analysis of a reference vehicle were also conducted during the first burning stage. The numerical results obtained with the present orthotropic model and the dynamic analysis method were found to be in good agreement with the natural vibration characteristics according to previous analyses and experiments. Additionally, the reference vehicle showed the estimated occurrence of pogo in the first structural mode when compared with the frequencies of the propellant feeding system. In conclusion, the present structural modelling and modal analysis procedures can be effectively used to analyse dynamic characteristics of liquid propellant launch vehicles. These techniques are also capable of identifying the occurrence of pogo and providing design criteria related to pogo instability.

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BARAUSKAS, Rimantas, Algimantas FEDARAVIČIUS i Karolis JASAS. "Structural Model Analysis for the Laser Guided Mobile Short-Range Air Defence System". Problems of Mechatronics Armament Aviation Safety Engineering 14, nr 1 (31.03.2023): 9–22. http://dx.doi.org/10.5604/01.3001.0016.2957.

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This article provides structural model analysis of characteristics of the vehicle movement during the shooting from a laser guided short-range air defence system mounted on it. Every shot causes a recoil, which significantly affects vibration of the vehicle and determines the shortest time duration between each two successive shots. The numerical simulation with MATLAB software enabled us verification of the created structural model and it facilitated the understanding of mechanical phenomena which are the most important for achieving the proper vibrational characteristics of the system. Main modes of the system movements during missile launch and their effect on the launcher and the vehicle were determined.

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Krivenko, Olga, i Petro Lizunov. "Vibrations of launch vehicle fairings with conical shape". Strength of Materials and Theory of Structures, nr 109 (11.11.2022): 66–71. http://dx.doi.org/10.32347/2410-2547.2022.109.66-71.

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Vibrations of launch vehicle conical fairings are investigated. Fairings are simulated using thin conic shells. The modal analysis of a thin elastic shell is based on the use of the developed finite element model of an inhomogeneous shell. In general, the technique makes it possible to investigate the geometrically nonlinear deformation, stability, and post-buckling behavior of a wide class of thin elastic shells. The modal analysis of the structure is implemented at each step of the static thermomechanical load. The subspace iteration method is used to determine the spectrum of the lowest vibration frequencies of shells of an inhomogeneous structure. The shell behavior analysis method is based on the relations of the three-dimensional theory of thermoelasticity and uses the finite element moment scheme. A thin elastic shell is simulated by a universal solid isoparametric finite element. The parameters of natural vibrations of conical shells of revolution with different thicknesses are investigated. Comparison of the calculation results obtained by the finite element moment scheme with the data of other authors shows a fairly good agreement between the solutions.

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Gong, Zheng, Zian Wang, Chengchuan Yang, Zhengxue Li, Mingzhe Dai i Chengxi Zhang. "Performance Analysis on the Small-Scale Reusable Launch Vehicle". Symmetry 14, nr 9 (6.09.2022): 1862. http://dx.doi.org/10.3390/sym14091862.

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According to the symmetrical characteristics of a new type of Reusable Launch Vehicle (RLV) in the recovery phase, we studied the basic aerodynamic model data of Starship and the aerodynamic data with rudder deflection, and the causes of its aerodynamic coefficients are expounded. At the same time, we analyzed its stability and maneuverability. According to the flying quality requirements, the lateral-directional model of Starship in the return phase at a high angle of attack is analyzed. Finally, we analyzed the lateral heading stability and control deviation of Starship by using the criterion and nonlinear open-loop simulations. The results show that the Starship has pitching and rolling stability, but it only has heading stability in some ranges of angle of attack, and there is no heading stability at a conventional large angle of attack. At the same time, after modal analysis and comparison of flight quality, it can be seen that the longitudinal long-period model of the starship degenerates into a real root and it is stable and convergent. The lateral heading roll mode is at level 2 flight quality, the helical mode is at level 1 flight quality, and the Dutch roll mode diverges, which needs to be stabilized and controlled later.

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Cole, Stanley R., i Thomas L. Henning. "Buffet response of a hammerhead launch vehicle wind-tunnel model". Journal of Spacecraft and Rockets 29, nr 3 (maj 1992): 379–85. http://dx.doi.org/10.2514/3.26362.

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Sarae, Wataru, Keita Terashima, Seiji Tsutsumi, Tetsuo Hiraiwa i Hiroaki Kobayashi. "Results of subscale model acoustic tests for H3 launch vehicle". Journal of the Acoustical Society of America 142, nr 4 (październik 2017): 2490. http://dx.doi.org/10.1121/1.5014087.

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Guo, Zhong Quan, Jian Xia Liu i Wen Cai Luo. "Parametric Modeling and Simulation for Aerodynamic Design of Launch Vehicle". Applied Mechanics and Materials 101-102 (wrzesień 2011): 697–701. http://dx.doi.org/10.4028/www.scientific.net/amm.101-102.697.

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Aerodynamic design of launch vehicle is facing combinatorial explosion problem caused by modular design. In order to get basic feasible solution from huge design space, the efficiency of design and simulation must be improved. In this paper, a parametric modeling and simulation method is proposed, which is based on CAD/CFD tools. Firstly, the design Variables of the launch vehicle are divided into three categories: size parameters, configuration parameters and mesh parameters. Secondly, parametric geometry model, including size and configuration parameters, is obtained by secondary development of Pro/ENGINEER. Thirdly, parametric mesh files for CFD are generated by implementing CFD-GEOM with scripts written in Python. By specifying boundary conditions through command stream of GAMBIT, FLUENT software will run automatically to calculate the aerodynamic performance of the launch vehicle. Finally, a graphical user interface (GUI) is developed using VC++6.0. With this system, the integration of CAD/CFD application is achieved. As long as designers enter certain design parameters in the GUI, they will quickly achieve 3D geometry model and aerodynamic performance of the launch vehicle. Application examples show that, this system can significantly improve the efficiency of aerodynamic design of the launch vehicle, and the data error between simulation and experiment is less than 10%, which is acceptable.

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Chen, Haipeng, Kang Chen i Wenxing Fu. "Memory Augmented Neural Network-Based Intelligent Adaptive Fault Tolerant Control for a Class of Launch Vehicles Using Second-Order Disturbance Observer". Mathematical Problems in Engineering 2021 (1.07.2021): 1–12. http://dx.doi.org/10.1155/2021/9961278.

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This paper focuses on the MANN-based intelligent adaptive fault tolerant control for a class of launch vehicles. Firstly, the attitude dynamic model of the launch vehicles suffering from the actuator faults and disturbances has been formulated. Secondly, the second-order disturbance observer has been designed for the launch vehicle to achieve the exact estimation and compensation of the time-varying disturbances. Meanwhile, the MANN has been introduced as online approximator, suppressing the adverse influence of the unknown nonlinearities. Moreover, several adaptive laws have been proposed to achieve the quick response to the actuator faults and the update of the MANN weights. As a result, the MANN-based intelligent adaptive fault tolerant control structure has been constructed for the launch vehicles. It has been proven that all the signals in the closed-loop system are bounded. Simulation results demonstrate the desired performance and the advantages of the proposed control algorithm.

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Hassani, Mehdi, Jafar Roshanian i A. Majid Khoshnood. "A reliable analytical navigation system based on symmetrical dynamic behavior of control channels". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, nr 1 (2.10.2016): 190–99. http://dx.doi.org/10.1177/0954410016664917.

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This study develops an advanced fault recovery strategy to improve the reliability of an aerospace launch vehicle navigation system. The proposed strategy contains fault detection features and can reconfigure the system against common faults in the aerospace launch vehicle navigation system. For this purpose, fault recovery system is constructed to detect and reconfigure normal navigation faults operating as a soft sensor, based on the symmetrical dynamic behavior of the yaw and pitch channels of the vehicle. In the face of pitch channel sensor failure, the Auto Regression Exogenous model of the yaw channel of the vehicle is identified using the recursive instrumental variable methodology. Based on the symmetrical behavior of the aerospace launch vehicle in the yaw and pitch channels, the Auto Regression Exogenous model of the yaw channel is substituted by the dynamic model of the pitch one and consequently, the pitch-rate gyroscope output is constructed to provide fault-tolerant navigation solution. The novel aspect of paper is employing symmetrical dynamic behavior of the yaw and pitch channels as an online tuning of analytical fault recovery solution against unforeseen variations due to its hardware/software property. In this regard, a nonlinear model of the aerospace launch vehicle is simulated using specific navigation failures and the results verified the feasibility of the proposed system. Simulation results and sensitivity analysis show that the proposed techniques can produce more effective estimation results than those of the previous techniques, against sensor failures.

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Wang, Tao, Ze Huan Huang, Hui Zhang i Bin Zheng. "The Displacement Measurement of Ground Wind Loads for Launch Vehicle". Applied Mechanics and Materials 568-570 (czerwiec 2014): 100–105. http://dx.doi.org/10.4028/www.scientific.net/amm.568-570.100.

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This paper presented a method of the displacement measurement of ground wind loads for launch vehicle based on laser and Digital Image Processing Technology. The launch vehicle structure model was built by calculating the wind pressure. Measurement model and algorithm for single laser were given. In view of nonrigid component, a method for computing the centroid displacement was presented. Error analysis and error reducing measures were preliminarily discussed.

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Ragul, MS, Vishnu Prakash, G. Arshiya i Ankit Kumar Mishra. "Theoretical Model Study on Chemical Compositions Affecting the Space Launch Vehicles". 1 8, nr 1 (1.02.2022): 35–38. http://dx.doi.org/10.46632/jemm/8/1/6.

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In this paper we clearly discussed about composition of chemical substances that affects the launch vehicle in ground station as well as atmospheric conditions. During the rocket-launching the large amount of inhalation of an exhaust gas released and it majorly affects the surface of the launch pad as well as the atmosphere. In rocket, the combustion produces huge number of hot gases with high temperature and pressure. This hot gas passes into the nozzle and accelerates, that time hot cloud is formed in the ground station which composed of (CO2) Carbon dioxide, (HCl) Hydrogen chloride, and carbon monoxide (CO). These hot gases not only affect the ground surface but also the atmospheric layers, will see how it’s affecting the launch vehicle as well as to protect the environment from air pollution affected by emission of gases from different types of rockets and spacecrafts

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Jin, Yinxiang, Dawei Wang, Zhangxia Guo, Libo Zou, Jiahao Chen i Jingyun Xie. "Research on Electromagnetic Protection Characteristics of Vehicle-mounted Electromagnetic Railgun Launcher Lifetime Counting Device". Journal of Physics: Conference Series 2478, nr 12 (1.06.2023): 122021. http://dx.doi.org/10.1088/1742-6596/2478/12/122021.

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Abstract When the vehicle-mounted electromagnetic gun is launched, the vehicle body is subjected to strong arc reaction, magnetic field interference and shock wave phenomenon. By establishing the external magnetic field model of the multi-tube launch chamber, the electromagnetic protection characteristics of the vehicle-mounted electromagnetic gun launcher’s full-life counting device are studied, and the magnetic field characteristics during launch are analyzed by COMSOL software, and the distribution position of the counting device is selected. Then, based on the electromagnetic shielding technology, conduct anti-interference technology research on the counting device, adopt the method of combined shielding, compare and analyze the shielding effectiveness of each shielding body, and propose the best shielding scheme. The life counting device works safely and normally. Therefore, the research method in this paper can provide an electromagnetic shielding scheme for the anti-jamming technology of the remaining electronic equipment of the vehicle-mounted electromagnetic gun, and can provide guidance for the design of the vehicle-mounted multi-barrel electromagnetic gun launching system.

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Zhao, Chen-geng, Zhong-yi Sun, Yi-fei Su, Yi-chen Wang i Gui-gao Le. "Study on bottom thermal environment of launch vehicle during high altitude flight". Journal of Physics: Conference Series 2364, nr 1 (1.11.2022): 012029. http://dx.doi.org/10.1088/1742-6596/2364/1/012029.

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Abstract Aiming at the problem of the thermal environment at the bottom of the launch vehicle in the high-altitude flight stage, taking the launch vehicle with core stage four-way machine as the model, the coupled convection / radiation heat transfer problem at the high-altitude of the launch vehicle is deeply studied through the numerical simulation method, the gas jet flow model of the launch vehicle in the high-altitude flight stage is established, and the flow field information of the rocket under different working conditions and the heat flow at the bottom of the rocket are obtained. The numerical results are compared with the experimental data to verify the effectiveness of the selected model and calculation method. The research shows that the external flow field of the rocket body will have a great impact on the thermal environment at the bottom. With the continuous increase of flight altitude, the convective heat flow at the bottom of the rocket body first rises and then decreases. In the high-altitude stage, the total heat flow at the bottom of the rocket body is dominated by radiant heat flow, the convective heat flow outside the main nozzle is dominant, and the radiant heat flow outside the traveling jet nozzle is dominant.

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Xue, Guo Hu, Jun Shan Mu, Hui Fen Li, Li Wei Zhu i Yang Liu. "Initial Orbit Determination Using Single Frequency GPS Measurements of Launch Vehicle". Applied Mechanics and Materials 599-601 (sierpień 2014): 964–69. http://dx.doi.org/10.4028/www.scientific.net/amm.599-601.964.

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Orbit determination by using GPS measurements of a launch vehicle is an important option for the initial orbit determination of the vehicle's payload, and its accuracy is higher than the results generated by radar measurements. However, only broadcast GPS ephemeris and clock products are used in the current GPS measurements processing method, and Klobuchar model is directly used. The paper proposes to use precise ephemeris and clock products, and adopts an improved ionospheric model based on altitude factor for GPS measurements processing. The dynamic smoothing method is further used. The numerical results show that the proposed method can improve the early orbit satellites orbit segment to determine accuracy.

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Xu, Erbao, Yan Li, Lining Peng, Yuxi Li i Mingshun Yang. "On-Line Interpretation and Real-Time Diagnosis of Rocket’s Single Equipment". Mathematical Problems in Engineering 2021 (12.03.2021): 1–12. http://dx.doi.org/10.1155/2021/6671403.

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The work state of a launch vehicle is generally interpreted automatically on software. However, the sheer number of target parameters makes it difficult to realize real-time interpretation, and abnormal interpretation result does not necessarily mean that the vehicle is in abnormal state. This paper introduces the edge computing to achieve on-line interpretation and real-time diagnosis of a single launch vehicle. Firstly, the parameters to be interpreted were subjected to thresholding, leaving only those with high interpretation value. Next, the interpretation server layer of the real-time diagnosis model was built based on the attribute and value reduction algorithm of variable precision rough set (VPRS). Moreover, the higher-grade criteria were written in criterion modeling language (CML) and used to interpret the various higher-grade interpretation data pushed by the edge layer in real time. On this basis, the outputs of the edge layer and interpretation server layer were integrated to achieve the real-time diagnosis of single vehicle faults. Finally, the proposed model was proved feasible through the application in a launch vehicle.

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Duraffourg, Elodie, Laurent Burlion i Tarek Ahmed-Ali. "Finite-time observer-based backstepping control of a flexible launch vehicle". Journal of Vibration and Control 24, nr 8 (1.09.2016): 1535–50. http://dx.doi.org/10.1177/1077546316664021.

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In this paper, a longitudinal model of a space launch vehicle was developed using the Lagrange mechanism and a free–free Euler–Bernoulli beam model. The aim was to propose a model including one flexible mode plus a nonlinear aerodynamic coefficient for nonlinear control design. We then studied the output feedback problem raised by using such a nonlinear model. The main achievement is to propose a new finite-time state observer when the measured outputs are corrupted by an unidentified flexible mode. This effect may destabilize a classical backstepping control law applied to the rigid model. To achieve this, a backstepping control law was redesigned to damp out the flexible mode, once measured and characterized. Hence a new adaptive finite time observer was developed. Closed-loop simulations show the effectiveness of the observer in combination with a redesigned backstepping control law when sensors and the launcher nozzle are collocated.

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37

Bakhtin, Aleksandr Georgievich, i Vasilii Aleksandrovich Titov. "VERIFICATION OF A COMPUTATIONAL DYNAMIC MODEL OF A LAUNCH VEHICLE STRUCTURE". TsAGI Science Journal 49, nr 1 (2018): 93–103. http://dx.doi.org/10.1615/tsagiscij.2018026788.

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Salton, Alexandria R., Michael M. James, Matthew F. Calton, Kent L. Gee, Reese D. Rasband, Daniel J. Novakovich i Brent O. Reichman. "Launch vehicle acoustic measurements for community noise model development and validation". Journal of the Acoustical Society of America 144, nr 3 (wrzesień 2018): 1673. http://dx.doi.org/10.1121/1.5067452.

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Park, Seoryong, Kiseop Yoon, Jeongwoo Ko, HanAhChim Choung, Seokjong Jang i Soogab Lee. "Integrated simulation model for prediction of acoustic environment of launch vehicle". Journal of the Acoustical Society of America 140, nr 4 (październik 2016): 3248. http://dx.doi.org/10.1121/1.4970276.

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Fedaravičius, Algimantas, Karolis Jasas, Rimvydas Gaidys i Kęstutis Pilkauskas. "Dynamics of the Missile Launch from the Very Short-Range Mobile Firing Unit". Shock and Vibration 2023 (1.04.2023): 1–10. http://dx.doi.org/10.1155/2023/3082704.

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In this article, the results of research of dynamical processes during the first stage of a missile launch from very short-range mobile firing unit (MFU) are presented. The determined laws of motion of the system’s components enabled determining characteristic motions that need to be reconstructed by the launcher simulator intended to be developed for personnel training. The half-car approach is applied for modelling the unit which consists of a vehicle and is attached to its missile launcher. The combined system is represented by a lumped-parameter model. The mathematical model is derived by applying the principle of Lagrange differential equations. Motion laws of the components of the system as its response to excitation load generated during missile launch were determined.

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Li, Jiaxin, Donghui Wang i Weihua Zhang. "Surrogate-Based Optimization Design for Air-Launched Vehicle Using Iterative Terminal Guidance". Aerospace 9, nr 6 (1.06.2022): 300. http://dx.doi.org/10.3390/aerospace9060300.

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In recent years, the penetration of low-cost air-launched vehicles for nano/micro satellites has significantly increased worldwide. Conceptual design and overall parameters optimization of the air-launched vehicle has become an exigent task. In the present research, a modified surrogate-based sequential approximate optimization (SAO) framework with multidisciplinary simulation is proposed for overall design and parameters optimization of a solid air-launched vehicle system. In order to reduce the large computation costs of time-consuming simulation, a local density-based radial basis function is applied to build the surrogate model. In addition, an improved particle swarm algorithm with adaptive control parameters is proposed to ensure the efficiency and reliability of the optimization method. According to the LauncherOne air-launched vehicle, the overall optimization design problem aims to improve payload capacity with the same lift-off mass. Reasonable constraints are imposed to ensure the orbit injection accuracy and stability of the launch vehicle. The influences of the vehicle configuration, optimization method, and terminal guidance are considered and compared for eight different cases. Finally, the effect on the speed of optimization convergence of employing a terminal guidance module is investigated. The payload capability of the optimized configurations increased by 27.52% and 23.35%, respectively. The final estimated results and analysis show the significant efficiency of the proposed method. These results emphasize the ability of SAO to optimize the parameters of an air-launched vehicle at a lower computation cost.

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Chelaru, Teodor Viorel, i Adrian Chelaru. "Mathematical Model and Performance Evaluation for a Small Orbital Launcher". Applied Mechanics and Materials 772 (lipiec 2015): 388–94. http://dx.doi.org/10.4028/www.scientific.net/amm.772.388.

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The purpose of this paper is to present some aspects regarding the computational model and simulation for three stage launch vehicle (LV) used to inject in orbit small size payload. The computational model consists in numerical simulation of LV evolution for imposed start conditions. The launcher model presented will be with six degrees of freedom (6DOF) and variable mass. The results analysed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize a small multi-stage launcher, end evaluate his performance using the developed model. From technical point of view, the paper is focused on ESA project “Study – concept, to achieve a Small Orbital Launcher through zonal cooperation - SOL

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Zhao, Liangyu, i Yi Jiang. "A study on launch site ground of vehicle-mounted missile based on elastic layer theory". Advances in Mechanical Engineering 10, nr 10 (październik 2018): 168781401880732. http://dx.doi.org/10.1177/1687814018807322.

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With the widespread use of the vehicle-mounted missile launching system, the research of launch site ground becomes very important. The loading of missile launch system under the supporting disk is similar to inverted bell, and its radius and size of the missile launching are much more than a car. So the ready-made equivalent modulus formula under the action of circular uniformly distributed loading is not suitable for the calculation of launch site ground of vehicle-mounted missile. Therefore, based on the elastic layered system and the Gauss–Legendre numerical calculation method, a new calculation method of equivalent modulus is proposed, as well as the regression formulation for calculation of equivalent modulus of foundation is also shown. Based on the above analysis, a launch site ground mechanical model is presented. The thin plate model on elastic half-space foundation was formulated taking into account normal direction and the double-layer elastic system model was proposed for in tangential direction. Validity of the model was done by coring the programming calculation with the finite element numerical simulation for selected cases.

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Jayaprasad, G., P. P. Dhanlakshmi i S. Hemachandran. "Analysis of electrical discontinuity problem in MLB using Ishikawa model". Circuit World 42, nr 4 (7.11.2016): 201–6. http://dx.doi.org/10.1108/cw-08-2016-0036.

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Purpose The purpose of this study/paper is regarding analysis of electrical discontinuity in penultimate layer of a few batches of Multilayer Boards (MLB) fabricated and supplied by a vendor. The ever-increasing demand of miniaturization in launch vehicle and spacecraft electronics systems has led to the usage of multilayer printed circuit boards (PCBs) for realizing high-performance electronics circuitry. Multilayer boards (MLBs) fabricated by qualified agencies based on the customer requirement are being used in the critical launch vehicle/spacecraft systems after evaluating the preliminary test results supplied by the vendor. However, a few batches of MLBs fabricated and supplied by a particular vendor (“A”) showed a discontinuity problem in a few PCB tracks connected by soldering pads. As these MLBs are part of Flight critical systems of both launch vehicle and spacecraft, a malfunction in the board may lead to fatal errors during fight or on-orbit, thereby jeopardizing the mission. Design/methodology/approach A systematic approach was followed to have a thorough understanding of the problem, and major tests such as inspection, continuity measurement, microsection of the plated through hole (PTH) and Scanning Electron Microscopy–Energy Dispersive X-ray Analysis tests were conducted on identified test boards based on Ishikawa model. Emphasis was given for horizontal microsection, as it has got a clear edge in detecting defects at any point of PTH barrel to inner-layer copper interface. Findings Systematic testing and evaluation on specimen revealed the presence of unwanted material at the bonding area of inner-layer copper and PTH copper due to inadequate fabrication process. The un-cleaned epoxy materials present at the bonding area creates a weak bond between barrel and inner-layer copper. Electrical strength of the MLB is the strength of this link. This weaker interconnection leads to electrical discontinuity of inner-layer tracks. Originality/value MLBs are part of Flight critical systems of both Launch Vehicle and Spacecraft; a malfunction in the board may lead to fatal errors during fight or on-orbit, thereby jeopardizing the mission. Case study of an original failure observed in MLBs helped to achieve normal functioning of systems and avoided failures at later stage of mission.

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Aprovitola, Andrea, Luigi Iuspa i Antonio Viviani. "Thermal Protection System Design of a Reusable Launch Vehicle Using Integral Soft Objects". International Journal of Aerospace Engineering 2019 (28.04.2019): 1–14. http://dx.doi.org/10.1155/2019/6069528.

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In the present paper, a modelling procedure of the thermal protection system designed for a conceptual Reusable Launch Vehicle is presented. A special parametric model, featuring a scalar field irradiated by a set of bidimensional soft objects, is developed and used to assign an almost arbitrary distribution of insulating materials over the vehicle surface. The model fully exploits the autoblending capability of soft objects and allows a rational distribution of thermal coating materials using a limited number of parameters. Applications to different conceptual vehicle configurations of an assigned thickness map, and material layout show the flexibility of the model. The model is finally integrated in the framework of a multidisciplinary analysis to perform a trajectory-based TPS sizing, subjected to fixed thermal constraints.

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Escartí-Guillem, Mara S., Luis M. García-Raffi i Sergio Hoyas. "URANS Analysis of a Launch Vehicle Aero-Acoustic Environment". Applied Sciences 12, nr 7 (25.03.2022): 3356. http://dx.doi.org/10.3390/app12073356.

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Predicting and mitigating acoustic levels become critical because of the harsh acoustic environment during space vehicle lift-off. This paper aimed to study the aero-acoustic environment during a rocket lift-off. The sound propagation within a launch event was studied using dedicated computational fluid dynamics (CFD). The resolution of all the phenomena that occur is unfeasible. We discuss the turbulence simplification and propose a feasible simulation through an unsteady Reynolds-averaged Navier–Stokes (URANS) model. The results were validated with experimental data showing a good correlation near the fairing surface and an improvable accuracy in the far field. To assess noise generation, the main shock waves were identified, and the evolution of the generated sound pressure was assessed. Moreover, vertical directivity was revealed by data analysis of the pressure field surrounding the fairing.

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Kim, Hong-Rae, Dong-Seo Yoo, Jong-Kwon Choi i Young-Keun Chang. "Cost Model for Annual Cost Spread Estimation of Space Launch Vehicle Development". Journal of the Korean Society for Aeronautical & Space Sciences 39, nr 6 (1.06.2011): 576–84. http://dx.doi.org/10.5139/jksas.2011.39.6.576.

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Krishnan, Ranjani, i V. R. Lalithambika. "Modeling and Validating Launch Vehicle Onboard Software Using the SPIN Model Checker". Journal of Aerospace Information Systems 17, nr 12 (grudzień 2020): 695–99. http://dx.doi.org/10.2514/1.i010876.

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Jia, Juhong, Debin Fu i Zepeng He. "Aerodynamic interactions of a Reusable Launch Vehicle model with different nose configurations". Acta Astronautica 177 (grudzień 2020): 58–65. http://dx.doi.org/10.1016/j.actaastro.2020.07.022.

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Sun, Wei, Yu Long Hua i Guo Qiang Liu. "A Test Study of Wet Dual Clutch Transmission during Vehicle Launch". Advanced Materials Research 490-495 (marzec 2012): 86–90. http://dx.doi.org/10.4028/www.scientific.net/amr.490-495.86.

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This paper focuses on the model and analysis of wet dual clutch transmission (DCT) during vehicle launch. Two evaluation indexes, slipping friction work and degree of jerk, is presented, and a single clutch control strategy is established and has been validated after applied for an experimental vehicle equipped with a wet dual clutch transmission.

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