Search

A three-dimensional mathematical model of the diffusion of impurities from a short-term source taking into account meteorological conditions is presented. The model includes the system of equations of hydrothermodynamics describing regional atmospheric processes. The transfer of multicomponent gas impurities is calculated taking into account microphysical processes, precipitation, and fog. Initial information for model initialization is aerological information in the form of actual or forecast fields of meteorological parameters and characteristics of sources of impurities. A numerical analysis of the diffusion of impurities in the near zone of a cosmodrome from short-term sources such as engines of launch vehicles was performed for various meteorological parameters, including wind in the atmosphere. [ABSTRACT FROM AUTHOR]

The importance of Fractional Order Proportional Integral Differential (FOPID) controller is increased due to the presence of two design parameters more than the conventional PID controller. This paper uses of these additional degrees of freedom in designing the attitude controller for launch vehicle system with unstable rigid body dynamics and flexible dynamics. The parameters of the controller are designed to meet time domain specifications by optimizing the performance indices using genetic algorithms. The proposed tracking controller exhibits the robustness property to the parameter variations during highly unstable atmospheric phase of the attitude controller. The tracking performance of the controller is analysed with flexible vehicle dynamics in presence of two bending modes. The proposed controller can reduce the steady state error by 30% and the settling time is reduced by 50%. The time domain analysis confirms the effectiveness of the proposed controller even in the presence of unmodelled dynamics and wind disturbance without losing stability of launch vehicle. In order to assure the robustness property of the proposed controller, the plant parameters such as control moment coefficient and aerodynamic coefficients are varied by 25%. The tracking performance with these off nominal parameters shows only 1% deviation from the response of plant with nominal parameters. [ABSTRACT FROM AUTHOR]

The Crew Escape System (CES) is used for safe return of crew in case of any malfunctioning of a launch vehicle during ascent, specially in the atmospheric phase of flight. In the present case, as no active control is employed, the aerodynamic stability of CES is essential for safe return of crew in case of any emergency during ascent phase. While finalizing the shape of CES, besides stability, other aerodynamic parameters like acoustic, thermal, structural load, etc. need to be analysed. Based on detailed studies using CFD, a conical-shaped CES has been finalized. This article discusses the challenges involved in aerodynamic shaping of CES and different parameters studied while finalizing the configuration. [ABSTRACT FROM AUTHOR]

Crew Escape System (CES) motors for Human Rated Launch Vehicle (HRLV) have been newly designed. These solid motors have special nozzle configuration. The development of these motors was a challenging task due to stringent design specification and intricate nozzle configuration. This paper emphasizes the fabrication challenges encountered during realization of CES hardware, details of process and quality improvement done, to overcome the challenges faced during fabrication and to achieve the stringent geometrical requirements and dimensional accuracies as specified by designer. [ABSTRACT FROM AUTHOR]

Abort system is initiated to take the Crew Module (CM) away from the launch vehicle in case of an emergency at lift-off or at any point of time after launch for a mission with crew onboard. Crew Escape System (CES)-based abort is carried out from launch pad and during the atmospheric phase of ascent flight. The design and operation of CES play a crucial role in providing abort capability for escape from launch vehicle and return of the crew back to Earth during critical phase of ascent flight. CES motors are used to pull CM away from the launch vehicle during this mode of abort. Mission simulation and analysis is necessary for the design of CES-based abort mission and for its configuration. This article discusses the mission design, challenges faced during the design and strategies formulated towards the successful execution of Pad Abort Test. [ABSTRACT FROM AUTHOR]

Coupled loads analysis (CLA) has been used in order to evaluate structural design of satellite for vibration environment during launch and micro-vibration disturbance environment on-orbit. Each time the design of satellite’s substructure is updated, CLA, which is to calculate coupled vibration responses by using coupled whole model, needs to be performed. In the case of frequently re-calculation, CLA is not an efficient means due to its costly calculation caused by coupling of updated models, defining forcing function and managing and assessment re-calculation result of finite element analysis. Therefore, there is a need for a method for efficiently calculating coupled vibration responses. In this paper we propose a quick update method of coupled vibration responses to evaluate updated design efficiently without the costly calculation. In the proposed method, the initial accelerations at the boundary of substructures without coupling of them are once calculated, and then the initial boundary accelerations can be updated to the coupled boundary accelerations by simple matrices calculation using Craig-Bampton model parameters of substructures. We show a calculation example of the proposed method with simulation models, which demonstrates that the coupled accelerations calculated by the proposed method correspond well with those of finite element analysis results.

In spite of great progress in the rocket technology and practical experience in the development and production of a number of rocket space systems (i.e., launch vehicles – LV), the task of optimizing the LV structure is still essential. Even in the fundamental work of well-known American scientists [1] in 1965, where the task of optimization of propellant distribution between the stages of a multi-stage launcher configuration was raised, an approximate solution was proposed. It was shown that in case of equal specific impulses of rocket engines of LV stages working successively, the equal distribution of characteristic velocities for each stage is the optimal one. Later, the same conclusion was made in the teaching aids published in Russia [2]. The treatise [3], published in the US in 2004, says there is no other solution found yet. This paper suggested a new approach in solving this and some other related problems. [ABSTRACT FROM AUTHOR]

During the launch mission of a launch vehicle (LV), localization of the LV is very important for ensuring flight safety. Usually, active localization systems such as the ground tracking radar systems and onboard navigation systems of the LV are responsible for acquiring the relevant location information for localization. However, if, in addition to the existing localization methods, passive localization using ground telemetry stations can be implemented, the reliability of the flight safety control for the LV can be improved. Time of arrival (TOA), time difference of arrival (TDOA), and angle of arrival (AOA) are typical localization techniques used for an emitting target. In this paper, we examine passive TDOA/AOA localization that uses two ground telemetry stations, and propose a technique to improve its localization accuracy, based on second-order polynomial regression and the removal of the estimated multipath delays in the TDOA/AOA results. By comparing the results of our proposed technique with the measured data from an on-board GPS used in the 3rd Korea Space Launch Vehicle-1 (KSLV-1) mission, our study confirms that the localization accuracy has been improved. [ABSTRACT FROM AUTHOR]

In this paper, a novel composite sandwich structure analysis of Launch Vehicle (LV) fairing is considered and proposed by a new Multidisciplinary Design Optimization (MDO) for a two-stage launch vehicle. Accordingly, “Multidisciplinary Design Optimization based on Independent Subspaces” (MDOIS) is employed using the “Fixed Point Iteration” (FPI) method to achieve the best convergence at system level (SL) to segregate the disciplines. Therefore, two proposed subspaces overcome difficulties of common mentioned MDO of LVs. Hence, the first subspace is a MDO which includes propulsion, aerodynamics, weight and trajectory disciplines and the second one, includes the novel composite fairing structure optimization as the other single discipline optimization that considered as a compact problem analytically and numerically and it is one of the novelties of this work. By considering variables as propulsion, trajectory and also composite sandwich fairing structure design regarding to the variables of designing and the performing optimization process, the fairing mass has been reduced more and considerable with respect to the common two stages LVs. In addition, due to the global optimization of LVs this weight reduction caused in reduction of the total gross weight of LVs. This system engineering proves the high sufficiency of MDO in complicated designing and it can be a roadmap for the future space vehicles designers especially who want to consider the composite structure optimization in LVs.

This paper solves the problem of the on-line joint optimization of the target orbit and flight trajectory of a launch vehicle experiencing a thrust drop failure, and proposes state-triggered indices to improve the solving efficiency. The 'highest-height index' is triggered first to find the maximum-height circular orbit (MCO) in the orbital plane formed at the time of failure. If the orbital altitude is less than the perigee height of the prescribed target orbit, the MCO will be regarded as the optimal circular orbit to ensure safety. If not, the 'minimal orbital plane deviations index' is triggered to subsequently find the optimal elliptical orbit (OEO), where the orbital inclination and longitude of the ascending node are preferentially adjusted on the basis of the MCO. Furthermore, if the OEO is coplanar with the target orbit, a 'minimal orbital shape deviations index' is triggered to find the optimal rescue orbit, which eliminates the deviations of the perigee altitude, semi-major axis and perigee argument on the basis of the OEO. Estimation of the geocentric angle of the injection point, construction of the approximated sub-problem, and convexification processing are adopted to provide initial guesses to boost the computing. Monte Carlo simulations show that the method not only converges to the optimal or sub-optimal solution for different failure states, which avoids suspended calculations, non-convergence, and constraint violations that are common when directly solving this type of rescue problem, but also meets the basic requirements of on-line applications. • State triggered index is proposed to get optimal rescue orbit if thrust drop occurs. • Orbital altitude, plane, and shape are adjusted sequentially to ensure solvability. • Convexification and range estimation are used to get initial value and boost solving. • Simulations show the adaptability to various fault scenarios and on-line application. [ABSTRACT FROM AUTHOR]

The article describes the vegetation cover that formed at three crash sites of launch vehicles fueled by toxic unsymmetric dimethylhydrazine 4–12 years after the accident. A launch-vehicle crash is accompanied by chemical, mechanical, and thermal (or pyrogenic) effects on the vegetation cover that extend out from the impact site to different distances, thereby forming characteristic, technogenically disturbed areas. The vegetation cover at the three 2006–2013 crash sites of launch vehicles are characterized by geobotanical descriptions that were performed in plots located at different distances from the center of the impact site. A self-restoration feature of the vegetation cover at the studied crash sites is the formation of an annual saltwort community (Climacoptera lanata, Salsola laricina, S. paulsenii, etc.) with a relatively high projective cover on the site of the former impact funnel due to soil subsidence and surface-runoff redistribution. The original vegetation regenerates with distance from the center of the impact site: the proportion of annual plants decreases, while the total projective cover and floristic saturation of the communities increases. According to the most optimistic forecasts, the period of self-restoration of phytocenoses without biological reclamation can be no less than 15 years. [ABSTRACT FROM AUTHOR]

The problem of optimal control of stochastic discrete parameter-varying uncertain linear systems is formulated and partially solved. The solution is achieved by a generalization of the State and Parameters Observability Canonical form - SPOC to parameter varying systems. This representation of linear time-variant systems enables the application of tools from the LQR-LQG theory of control and estimation of discrete linear parameter-varying systems. The optimal solution is exact and noncausal. A causal suboptimal controller, using certainty equivalence, is proposed as an adhoc solution. This controller needs only the knowledge of the order of the system. The scheme is BIBO stable for sufficiently low noises and sufficiently slowly varying parameters. As an example, the proposed algorithm is applied to a difficult problem: control of an unstable nonminimum phase parameter-varying model of a dynamic launch vehicle with very large uncertainty, and unmodelled dynamics. Care is devoted to analysis of the robustness of the control algorithm. It is demonstrated via simulations that as long as the assumptions are valid the algorithm remains stable and preserves performance. [ABSTRACT FROM AUTHOR]

This manuscript presents experimental results of compression tests of repaired sandwich structure typical of a launch vehicle Payload Adapter Fitting (PAF) that is being considered for the National Aeronautics and Space Administration's (NASA) Space Launch System (SLS) launch vehicle. Specimens with 8-ply quasi-isotropic carbon/epoxy face sheets and aluminum honeycomb core were damaged with representations of misdrilled holes, barely visible impact damage (BVID) or visible impact damage (VID). A patch repair technique not requiring removal of any material is presented and assessed for residual in-plane compressive load-carrying capability on each of the three types of damage. Edgewise compression tests on specimens demonstrated the repairs could recover all measured undamaged compressive load-carrying capability of the misdrilled hole and BVID specimens. For the specimens with VID, an average of about 91% of the average measured undamaged compressive load carrying capability was recovered indicating that larger patches, or perhaps removal and replacement of the damaged material, may be needed for more severe damage if all the measured undamaged compression strength needs to be regained. [ABSTRACT FROM AUTHOR]

The German Aerospace Center (DLR) is currently developing the Reusability Flight Experiment (ReFEx). The successor of the already performed hypersonic flight experiments SHEFEX I and II shall be launched on a Brazilian VSB-30 sounding rocket in 2022 and shall achieve a re-entry velocity of more than Mach 5. The main goals of the project are the demonstration of a controlled autonomous re-entry flight from hypersonic velocity down to subsonic range and the testing of the key technologies required for future reusable first stage systems. Utilizing Concurrent Engineering (CE) approach the fundamental feasibility of this sophisticated flight experiment has been investigated by the entire ReFEx team. All required systems, including sensors and actuators as well as their interfaces have been defined and different options were assessed regarding matters such as the scientific output, complexity, risk and cost. The current configuration of ReFEx has a re-entry mass of about 400 kg, a length of 2.7 m and a wingspan of 1.1 m. This paper provides a system overview and addresses the Reusable Launch Vehicles (RLV) technologies Guidance, Navigation and Control as well as Flight Instrumentation. Furthermore the mission design (launch & re-entry) and the main challenges regarding the mission realization are addressed. • Perform a controlled flight in the velocity range hypersonic down to subsonic. • Perform a controlled heading change. • Development of key technologies for winged RLV first stages. • Perform In Flight Data acquisition using advanced sensors for RLV research and development purposes. • Recovery of the experimental vehicle. [ABSTRACT FROM AUTHOR]

An efficient and accurate method was proposed for dynamic characteristics prediction of launch vehicles based on the equivalent beam model, with which the influence of liquid propellants can be considered. The equivalent beam model for the launch vehicle was established based on the cross-sectional area equivalence principle, and improved through model updating against the natural frequencies and mode shapes of the fine finite element model. The present method was validated through computation of a certain type of launch vehicles. A highly representative equivalent beam model for the launch vehicle was established. The dynamic characteristics of the launch vehicle were predicted based on the equivalent beam model under the influence of liquid propellants in 2 ways of the lumped mass method and the coupled mass method, and the results were compared. [ABSTRACT FROM AUTHOR]

In the article, the analysis of the provisions of the Missile Technology Control Regime enables to conclude that it requires to be amended, or, which is better, replaced by a new high-level international legal treaty. The authors reveal that classic civilian launch vehicles built by private space companies have a common good for humanity; in particular, they increase the chance of human survival as a result of global natural or human-made disasters. The authors conclude that a new paradigm should be implemented: from restraining the transfer of rocket technology to encouraging this process by space companies under the banner of democracies governed by the rule of law. At the same time, the authors argue that more effective international legal sanctions and controls should be implemented by the international community to prevent ballistic missiles from getting to terrorists. [ABSTRACT FROM AUTHOR]

The article studies the current state of affairs in national space activities. The legal problems facing the space industry at the present stage, the mechanism of their solution are analyzed. The author highlights the crucial issue, which requires an immediate solution regarding the location of the launch site for domestic space objects. The special role of the State in ensuring the legal conditions for the development of national space activities, which should be carried out using good in law modern methods and means of public regulation, is determined. Nowadays, the legislation of Ukraine on space activities requires significant changes and additions, taking into account the new space realities and instruments for the development of public relations in the field of space exploration and use. The author argues that a strategy for the development of space activities in Ukraine, public administration of the space industry, as well as ensuring its restructuring require a substantial rethinking and improvement. The relevance of the legal framework for the creation and use of space defense and dual-use space technology in the interests of the national security and defense sector is underlined. The author argues that the Concept of Ukrainian law reform on space activities, including the legal rationale for choosing the optimal behavior model regarding the solution of complex legal problems based on modern space-legal doctrine and practice of space relations, should be developed. [ABSTRACT FROM AUTHOR]