Gotowa bibliografia na temat „Launch Vehicle Model”

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.

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.

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.

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.

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.

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.

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.

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.

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.