Journal articles: 'Spacecraft Conceptual Design'

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Weber, A., S. Fasoulas, and K. Wolf. "Conceptual interplanetary space mission design using multi-objective evolutionary optimization and design grammars." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 225, no. 11 (September 9, 2011): 1253–61. http://dx.doi.org/10.1177/0954410011407421.

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Conceptual design optimization (CDO) is a technique proposed for the structured evaluation of different design concepts. Design grammars provide a flexible modular modelling architecture. The model is generated by a compiler based on predefined components and rules. The rules describe the composition of the model; thus, different models can be optimized by the CDO in one run. This allows considering a mission design including the mission analysis and the system design. The combination of a CDO approach with a model based on design grammars is shown for the concept study of a near-Earth asteroid mission. The mission objective is to investigate two asteroids of different kinds. The CDO reveals that a mission concept using two identical spacecrafts flying to one target each is better than a mission concept with one spacecraft flying to two asteroids consecutively.

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Mosher, Todd. "Conceptual Spacecraft Design Using a Genetic Algorithm Trade Selection Process." Journal of Aircraft 36, no. 1 (January 1999): 200–208. http://dx.doi.org/10.2514/2.2426.

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Kurenkov, Vladimir, and Alexander Kucherov. "Experiences in engineering design training at Samara University." SHS Web of Conferences 137 (2022): 01013. http://dx.doi.org/10.1051/shsconf/202213701013.

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This paper aims to share experience gained by Samara National Research University in training students in engineering degree programme “Manned and Unmanned Spacecraft and Space Systems”. Core items of engineering degree curricula are projects because they basically form practical experience of future design engineers. Main items of the course project “Calculation of main parameters and generation of land remote sensing satellites conceptual design based on regard to required efficiency indices” are discussed. These include acquisition and processing of statistical data on space systems and satellites; determination of satellite orbit parameters; determination of massdimensional characteristics of spacecraft on-board systems and construction; formation of on-board systems; spacecraft construction models and assembly model; choice of launcher and spacehead model development. The activity is supported by appropriate training materials and software.

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Mahmoudi, M., A. B. Novinzadeh, and F. Pazooki. "Optimum conceptual design for the life support systems of manned spacecraft." Cogent Engineering 7, no. 1 (January 1, 2020): 1863304. http://dx.doi.org/10.1080/23311916.2020.1863304.

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Cuerno-Rejado, C., J. López-Díez, and A. Sanz-Andrés. "Rapid Method for Spacecraft Sizing." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 209, no. 3 (July 1995): 165–69. http://dx.doi.org/10.1243/pime_proc_1995_209_286_02.

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In this paper, a rapid method for spacecraft sizing is presented. This method is useful in both the conceptual and preliminary design phases of scientific and communication satellites. The aim of this method is to provide a sizing procedure similar to the ones used in the design of aircraft; actually by determining the mass of all the spacecraft subsystems. In the Introduction, the importance of an accurate initial mass budget in the design of satellites is emphasized. Literature about this topic is not very extensive and most of the existing methods have been recapitulated. The methodology followed in the proposed procedure for spacecraft mass sizing is based on these methods. Data from 26 existing satellites have been considered to obtain correlations between each subsystem mass and the mass of the whole spacecraft.

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Kerslake, Thomas W. "Effect of Voltage Level on Power System Design for Solar Electric Propulsion Missions." Journal of Solar Energy Engineering 126, no. 3 (July 19, 2004): 936–44. http://dx.doi.org/10.1115/1.1710523.

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This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and design points were generated for several system voltages using state-of-the-art or advanced technologies. A 300-V “direct-drive” architecture was also analyzed to assess the benefits of directly powering the electric thruster from the photovoltaic array without up-conversion. Computational models were exercised to predict the performance and size power system components to meet spacecraft mission requirements. Pertinent space environments were calculated for the mission trajectory and an electron current collection model was developed to estimate photovoltaic array losses due to natural and induced plasma environments. The secondary benefits of power system mass savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.

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Polites, Michael E., John P. Sharkey, Gerald S. Nurre, Philip C. Calhoun, and William D. Lightsey. "Advanced X-ray Astrophysics Facility-Spectrometry Spacecraft Pointing Control System - Conceptual design." Journal of Spacecraft and Rockets 32, no. 2 (March 1995): 344–52. http://dx.doi.org/10.2514/3.26616.

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Markley, F. L., F. H. Bauer, J. J. Deily, and M. D. Femiano. "Attitude control system conceptual design for Geostationary Operational Environmental Satellite spacecraft series." Journal of Guidance, Control, and Dynamics 18, no. 2 (March 1995): 247–55. http://dx.doi.org/10.2514/3.21377.

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Nagasaki, Y., T. Nakamura, I. Funaki, Y. Ashida, and H. Yamakawa. "Conceptual Design of YBCO Coil With Large Magnetic Moment for Magnetic Sail Spacecraft." IEEE Transactions on Applied Superconductivity 23, no. 3 (June 2013): 4603405. http://dx.doi.org/10.1109/tasc.2013.2243791.

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De Souza, Ariana C. Caetano, and Walter A. Dos Santos. "11.1.2 SpaceESB - A Proposal of an Enterprise Service Bus for Spacecraft Conceptual Design." INCOSE International Symposium 21, no. 1 (June 2011): 1272–80. http://dx.doi.org/10.1002/j.2334-5837.2011.tb01284.x.

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Kabganian, Masoud, Seyed M. Hashemi, and Jafar Roshanian. "Multidisciplinary Design Optimization of a Re-Entry Spacecraft via Radau Pseudospectral Method." Applied Mechanics 3, no. 4 (September 26, 2022): 1176–89. http://dx.doi.org/10.3390/applmech3040067.

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The design and optimization of re-entry spacecraft or its subsystems is a multidisciplinary or multiobjective optimization problem by nature. Multidisciplinary design optimization (MDO) focuses on using numerical optimization in designing systems with several subsystems or disciplines that have interactions and independent actions. In the present paper, the system-level optimizer, trajectory, geometry and shape, aerodynamics, and aerothermodynamics differential equations, are converted to algebraic equations using the Radau pseudospectral method (RPM) since a spacecraft is a nonlinear, extensive, and sparse system. The solution to the problem with the help of MDO is reached by iterating all the disciplines together; one can simultaneously enhance the design, decrease the time and cost of the entire design cycle, and minimize the structural mass of a re-entry spacecraft. Considering various methods presented in earlier research works, a combined and innovative all-at-once (AAO), RPM-based MDO method, including the key subsystems in the design process of a re-entry capsule-shape spacecraft with a low lift-to-drag ratio (L/D), is presented. Considering the applicable state and control variables, various constraints, and parameters applied to several geometric shapes of a blunt capsule and using Apollo’s aerodynamic and aerothermodynamic coefficients, the optimized dimensions for a re-entry spacecraft are presented. The introduced optimization scheme led to a 17% mass reduction compared to the original mass of the Apollo vehicle. Fast computing and simplified models are used together in this method to analyze a wide range of vehicle shapes and entry types during conceptual design.

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Kulkov, V. M., S. O. Firsyuk, A. M. Yurov, S. A. Tuzikov, Yu G. Egorov, and S. W. Yoon. "Principles of construction and application areas of small spacecraft based on unified space platforms." Spacecrafts & Technologies 6, no. 2 (June 24, 2022): 133–43. http://dx.doi.org/10.26732/j.st.2022.2.08.

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The conceptual options for small spacecraft on the basis of a unified space platform are analyzed, the areas of rational use of the unified space platform as part of the small spacecraft for solving a wide range of tasks, including monitoring of the Arctic regions, are explored. Methodical approaches are given to ensure the design analysis of circuit solutions of the unified space platform for small spacecraft remote sensing. The actual problem of selecting the standard size of the unified space platform, evaluating the effectiveness of its use, taking into account the necessary modernization for a specific payload, is being investigated. A technique and models for selecting the parameters for the modification of the remote sensing satellites have been developed. The application of the methodology makes it possible to determine the rational parameters for modifying the remote sensing spacecraft, to evaluate the impact of the features of design solutions and external relations on the total costs in the implementation of the project for the creation of the unified space platform. The relevance of research is related to the need to create competitive variants of small-sized spacecraft based on a unified space platform, designed to provide monitoring of high-latitude regions using the small spacecraft system.

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Govindarajan, Nivedha, Abhinav Jayaswal, Katari Gaurav, and Aravind Seeni. "Higher-Fidelity CFD Tools for Conceptual Design of Supersonic Business Jets." ECS Transactions 107, no. 1 (April 24, 2022): 515–32. http://dx.doi.org/10.1149/10701.0515ecst.

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Supersonic business jets are regarded as prime modes of next-generation supersonic transport. They allow long-distance travel at low duration by crossing the Mach 1 speed. This results in the breaking of the sound barrier and a sonic boom generated. Sonic boom and aerodynamic efficiency are the biggest concerns for the return of civil supersonic transport. Supersonic vehicles, such as spacecraft, high-speed missiles, and supersonic jets, are subjected to heavy shock wave drag during flight, which will seriously affect the aerodynamic performance of supersonic vehicles. In addition, supersonic vehicles are also subjected to heavy aerodynamic heating during flight. Therefore, modified geometries using aerospikes and aerodisks can be effective to reduce the above-said concerns. In this paper, we discuss aerospikes and aerodisks, its effectiveness, and its importance. The use of these novel mechanisms in the design of supersonic business jets is probed in order to perform future design studies.

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Alvarez, Jose. "Determination of Mass Properties for the VX-SAT Communication Satellite." Applied Mechanics and Materials 110-116 (October 2011): 415–22. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.415.

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The determination of a spacecraft’s mass properties is critical during its conceptual design phase. Obtaining reliable mass property information, early in the design of a spacecraft, could prevent design mistakes that can be extremely costly in the development process. The process of the determination of the mass properties consist in a systematic work, that include the taking part of all the sub systems or design teams that conforms the preliminary tasks for the design of the communications satellite. The VX-SAT project is a system engineering exercise, realized in the China Academy of Space Technology (CAST) like training for the future designs in communication satellites for the Bolivarian Agency of Space Activities (ABAE) in Venezuela. This paper aim to show an analysis of the procedures in the calculation of the mass properties in communication satellite using suitable software’s and the importance of the process to collect the mass budget information to get the more reliable results. Also includes an overview of the VX-SAT system, the preliminary design of the structure requirements and finally the methods used to the analysis of the general determination of the mass properties with its results.

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Fazeley, H. R., H. Taei, H. Naseh, and M. Mirshams. "A multi-objective, multidisciplinary design optimization methodology for the conceptual design of a spacecraft bi-propellant propulsion system." Structural and Multidisciplinary Optimization 53, no. 1 (August 22, 2015): 145–60. http://dx.doi.org/10.1007/s00158-015-1304-2.

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Talalaev, Aleksandr, Vitaly Fralenko, and Vyacheslav Khachumov. "Review of standards and the conceptual design of tools for spacecraft monitoring, control and diagnostics." Program Systems: Theory and Applications 6, no. 3 (2015): 21–43. http://dx.doi.org/10.25209/2079-3316-2015-6-3-21-43.

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Mateo-Velez, Jean-Charles, Jean-François Roussel, David Rodgers, Alain Hilgers, Marc Sevoz, and Patrice Pelissou. "Conceptual Design and Assessment of an Electrostatic Discharge and Flashover Detector on Spacecraft Solar Panels." IEEE Transactions on Plasma Science 40, no. 2 (February 2012): 246–53. http://dx.doi.org/10.1109/tps.2011.2173955.

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Viviani, Antonio, Luigi Iuspa, and Andrea Aprovitola. "Multi-objective optimization for re-entry spacecraft conceptual design using a free-form shape generator." Aerospace Science and Technology 71 (December 2017): 312–24. http://dx.doi.org/10.1016/j.ast.2017.09.030.

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Jeong, Ju Won, Young Ik Yoo, Jung Ju Lee, Jae Hyuk Lim, Kyung Won Kim, and Do Soon Hwang. "Development of a Novel Deployment Hinge Mechanism for a Spacecraft Using Axiomatic Design." Advanced Materials Research 314-316 (August 2011): 863–71. http://dx.doi.org/10.4028/www.scientific.net/amr.314-316.863.

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This paper presents a novel hinge mechanism for deployment of spacecraft subsystems such as antennas, solar arrays. By using Axiomatic design theory, the conceptual design of the hinge mechanism is suggested, which has not only high deployed stiffness and low deployment shock but also does not require lubrication and accurate fabrication. That is, optimization of deployment torque and maximization of the deployed stiffness can be possible since this suggested hinge mechanism is decoupled design. And the suggested hinge mechanism is fabricated and tested to evaluate the feasibility. Quasi-static analysis is performed to optimize deployment torque for low deployment shock by using FEM. Also, the bending stiffness is measured by 4 point bending test.

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Palii, O. S., E. O. Lapkhanov, and D. S. Svorobin. "Model of distributed space power system motion control." Technical mechanics 2022, no. 4 (December 15, 2022): 35–50. http://dx.doi.org/10.15407/itm2022.04.035.

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The goal of this article is to develop a generalized mathematical model for controlling the motion of the spacecraft of a space industrial platform’s distributed power system. Space industrialization is one of the promising lines of industrial development in the world. The development of space industrial technologies will allow one to solve a number of problems in the production of unique products unavailable under terrestrial conditions. The main types of these products include semiconductor materials, materials made by 3D printing in microgravity, space modules of sunshade systems, space metallurgy products, space debris processing products, and high-purity space biology substances. Taking this into account, a certain amount of electricity is required for the manufacture of one or another product. Given that some space industrial processes can consume a significant amount of electricity, a space industrial platform's own power generation may not be sufficient. Because of this, it was proposed to use additional energy resources through the development of a distributed power supply system for a space industrial platform. A group of power spacecraft is envisaged to collect and accumulate electric energy and transmit it in a contactless way to the receivers of the space industrial platform. The article presents mathematical models for the analysis of the orbital, angular, and relative motion of power spacecraft and receiver spacecraft. Algorithms are proposed for calculating the parameters of the power spacecraft orientation and stabilization system. A generalized model is constructed for determining the maximum distance and time interval of power spacecraft to platform electric power transmission using microwave radiation. The model developed allows one to choose the power spacecraft design parameters at the stage of conceptual design of space industrial platform power systems.

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Todorov, Vladislav T., Dmitry Rakov, and Andreas Bardenhagen. "Enhancement Opportunities for Conceptual Design in Aerospace Based on the Advanced Morphological Approach." Aerospace 9, no. 2 (February 1, 2022): 78. http://dx.doi.org/10.3390/aerospace9020078.

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The current challenges facing the aerospace domain require unconventional solutions, which could be sought in new configurations of future aircraft and spacecraft. The choice of optimal concepts requires the consideration of a significant amount of competing engineering solutions and takes place under conditions of uncertainty. Such a problem can be addressed by enhancing existing methods for analysis and synthesis solutions, such as the Advanced Morphological Approach (AMA). It uses morphological analysis to provide a more exhaustive overview of possible problem solutions, relies on expert evaluations of alternative technological options and applies clustering to the solution space. Although an intuitive method for structured concept generation, the AMA exposes the need for more robust problem structuring, improved objectivity of options evaluation and accounting for uncertainties. The current article suggests ways to overcome these challenges and their possible integration in the process. In particular, the integration of fuzzy sets is proposed to model uncertainties during the evaluation of technological options by the experts. The Fuzzy Analytical Hierarchy Process is adapted for integration into the AMA and for the conceptual design of aerospace vehicles.

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Palii, O. S. "Model for assessing the mass of a space industrial platform and its modules." Technical mechanics 2022, no. 3 (October 3, 2022): 75–84. http://dx.doi.org/10.15407/itm2022.03.075.

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The goal of this paper is to develop mass models of a space industrial platform and its modules. At the initial stage of development of a new spacecraft, a limited set of basic data is available. For a space industrial platform, they are as follows: the configuration of its main and auxiliary modules, the parameters of the technological processes to be implemented on the platform (the vacuum and the microgravity level, the equipment energy capacity), and the manufacturing equipment configuration. A feature of industrial platform design is that there are few, if any, theoretical works on the choice of platform parameters and the logic of platform conceptual design. In this paper, the design process is considered as applied to the conceptual design stage. This stage is characterized by that nothing is known about the system to be developed except for the general concept of the platform layout, the expected types of the main service systems, some basic data, and the parameters of the technological processes to be implemented on the platform. The process of designing a new complex space system such as an industrial platform is a multilevel iterative and optimization process, during which its characteristics and the mass fractions of its components are determined and refined. The paper presents a mass model of an industrial platform and its modules, in whose development the platform and its components were decomposed to the level of system elements. A statistical analysis of the mass fractions of the onboard spacecraft systems was carried out. The mean values of the mass fractions for the sample of spacecraft under study and their scattering coefficients (the dispersion and the mean square deviation) were determined. For the mean values and the dispersion, 99.9 confidence intervals were determined. Further studies on the design of space industrial platforms are planned to be carried using the mass fractions of satellite systems and the confidence intervals, namely, the minimum and the maximum possible mass for a particular system, determined in this study.

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Kim, Jae-Hoon, Kyun Ho Lee, Sung Kyung Hong, and Hae-Dong Kim. "Conceptual Design of Cold Gas Propulsion System of a Ground Simulator for Maneuver and Attitude Control Design Verification of Spacecraft." Journal of the Korean Society of Propulsion Engineers 19, no. 1 (February 1, 2015): 98–110. http://dx.doi.org/10.6108/kspe.2015.19.1.098.

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Deng, Li, Zhen Yang, Pengyu Du, and You Song. "A cloud platform for space science mission concurrent design." Concurrent Engineering 26, no. 1 (August 11, 2017): 104–16. http://dx.doi.org/10.1177/1063293x17724848.

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Using concurrent design methodology, the duration of space science mission in conceptual design phase can be shortened. This approach requires hardware and software resources to support, such as professional design tools and concurrent design environment. Except for the professional groups, the general researchers such as teachers and students have no chance to access these resources. Nowadays, more and more researchers distributed in different locations join into the space science research. The need for an open concurrent design platform offering design tools and data sharing environment has increased. This article presents a Cloud Platform of Concurrent Design for Space Science Mission. This Cloud Platform of Concurrent Design for Space Science Mission uses the idea of Software as a service, in which five design and analysis tools are offered as services to satisfy the basic requirements for space science mission concurrent design in conceptual design phase. This cloud platform provides the access to space science mission concurrent design for expert and non-expert users using a thin client of web browser. This article presents the platform architecture of the Cloud Platform of Concurrent Design for Space Science Mission and five software offered as services. The services include spacecraft conceptual orbit design, structure design, payload coverage analysis, data transmission analysis, and virtual community. And the basic cloud service (computing and storage) is also briefly introduced. It is described in detail how these services can be leveraged by users to do the concurrent design for one space science mission.

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Lee, Dukhang, Regina S. K. Lee, John E. Moores, Thomas Young, and Hugh Podmore. "Conceptual thermal design of a network of solar-powered Boardsat- and CubeSat-based landed spacecraft on Mars." International Journal of Space Science and Engineering 6, no. 2 (2020): 125. http://dx.doi.org/10.1504/ijspacese.2020.10032472.

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Lee, Dukhang, Thomas Young, Hugh Podmore, John E. Moores, and Regina S. K. Lee. "Conceptual thermal design of a network of solar-powered Boardsat- and CubeSat-based landed spacecraft on Mars." International Journal of Space Science and Engineering 6, no. 2 (2020): 125. http://dx.doi.org/10.1504/ijspacese.2020.110359.

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Pellegrino, Alice, Maria Giulia Pancalli, Andrea Gianfermo, Paolo Marzioli, Federico Curianò, Federica Angeletti, Fabrizio Piergentili, and Fabio Santoni. "HORUS: Multispectral and Multiangle CubeSat Mission Targeting Sub-Kilometer Remote Sensing Applications." Remote Sensing 13, no. 12 (June 19, 2021): 2399. http://dx.doi.org/10.3390/rs13122399.

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This paper presents the HORUS mission, aimed at multispectral and multiangle (nadir and off-nadir) planetary optical observation, using Commercial Off-The-Shelf (COTS) instruments on-board a 6-Unit CubeSat. The collected data are characterized by a sub-kilometer resolution, useful for different applications for environmental monitoring, atmospheric characterization, and ocean studies. Latest advancements in electro-optical instrumentation permit to consider an optimized instrument able to fit in a small volume, in principle without significant reduction in the achievable performances with respect to typical large-spacecraft implementations. CubeSat-based platforms ensure high flexibility, with fast and simple components’ integration, and may be used as stand-alone system or in synergy with larger missions, for example to improve revisit time. The mission rationale, its main objectives and scientific background, including the combination of off-nadir potential continuous multiangle coverage in a full perspective and related observation bands are provided. The observation system conceptual design and its installation on-board a 6U CubeSat bus, together with the spacecraft subsystems are discussed, assessing the feasibility of the mission and its suitability as a building block for a multiplatform distributed system.

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Russkikh, A. S., and V. V. Salmin. "Methodology of designing a space transport system including a “DM” chemical upper stage and an electric rocket transport module." VESTNIK of Samara University. Aerospace and Mechanical Engineering 21, no. 4 (January 18, 2023): 66–75. http://dx.doi.org/10.18287/2541-7533-2022-21-4-66-75.

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The article is devoted to the development of a methodology for designing a space transport system that has, in comparison with traditional injection scenarios, higher characteristics of the mass of the payload to be launched into high-energy orbits, including a geostationary one. The method is implemented by the systematization of design calculations and proposals for the rational design of interorbital vehicles with a combined propulsion system and the development of an algorithm for creating an electronic model of an interorbital vehicle in an automated designing PTC Creo system. The article considers algorithms for calculating the design parameters of the electric rocket transport module and ballistic parameters of the spacecraft flight to the target orbit. A method of forming conceptual design of an electric rocket transport module based on the obtained design parameters is also proposed. An electronic model of an electric rocket module was developed in a three-dimensional modeling system. The model displays the structural design of the module in accordance with the design parameters obtained and satisfies the design requirements imposed by other components of the space transport system. The proposed method makes it possible to design a space transport system that includes a DM upper stage and an electric rocket transport module designed to implement a combined payload injection scenario.

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Fusaro, Roberta, Nicole Viola, Franco Fenoglio, and Francesco Santoro. "Conceptual design of a crewed reusable space transportation system aimed at parabolic flights: stakeholder analysis, mission concept selection, and spacecraft architecture definition." CEAS Space Journal 9, no. 1 (June 27, 2016): 5–34. http://dx.doi.org/10.1007/s12567-016-0131-7.

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Ray, Asok, Min-Kuang Wu, Marc Carpino, and Carl F. Lorenzo. "Damage-Mitigating Control of Mechanical Systems: Part I—Conceptual Development and Model Formulation." Journal of Dynamic Systems, Measurement, and Control 116, no. 3 (September 1, 1994): 437–47. http://dx.doi.org/10.1115/1.2899239.

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A major goal in the control of complex mechanical systems such as advanced aircraft, spacecraft, and power plants is to achieve high performance with increased reliability, availability, component durability, and maintainability. The current state-of-the-art of control systems synthesis focuses on improving performance and diagnostic capabilities under constraints that often do not adequately represent the dynamic properties of the materials. The reason is that the traditional design is based upon the assumption of conventional materials with invariant characteristics. In view of high performance requirements and availability of improved materials, the lack of appropriate knowledge about the properties of these materials will lead to either less than achievable performance due to overly conservative design, or over-straining of the structure leading to unexpected failures and drastic reduction of the service life. The key idea of the research reported in this paper is that a significant improvement in service life can be achieved by a small reduction in the system dynamic performance. This requires augmentation of the current system-theoretic techniques for synthesis of decision and control laws with governing equations and inequality constraints that would model the properties of the materials for the purpose of damage representation and failure prognosis. The major challenge in this research is to characterize the damage generation process in a continuous-time setting, and then utilize this information for synthesizing algorithms of robust control, diagnostics, and risk assessment in complex mechanical systems. Damage mitigation for control of mechanical systems is reported in the two-part paper. The concept of damage mitigation is introduced and a continuous-time model of fatigue damage dynamics is formulated in this paper which is the first part. The second part which is a companion paper presents the synthesis of the open-loop control policy and the results of simulation experiments for transient operations of a reusable rocket engine.

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Averkiev, N. F., A. V. Kulvits, and T. A. Zhitnikov. "MULTILEVEL BALLISTIC STRUCTURE OF THE CLUSTER ORBITAL GROUPING OF REMOTE SENSING OF THE EARTH." Izvestiya of Samara Scientific Center of the Russian Academy of Sciences 23, no. 4 (2021): 77–85. http://dx.doi.org/10.37313/1990-5378-2021-23-4-77-85.

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The features of the application and justification of the orbital groupings of remote sensing of the Earth, consisting of clusters of small spacecraft, are considered. A review and analysis of the ballistic justification, construction and features of the use of orbital groups of remote sensing of the Earth is carried out. Modern approaches to the ballistic design of periodic review orbital groupings are considered. The article considers a new integrated approach to the ballistic construction of promising cluster orbital groupings, which will allow providing the spatio-temporal and accuracy characteristics required by the consumer, due to the optimal multi-level ballistic structure. The fundamental principles of constructing a cluster orbital grouping with a multi-level ballistic structure are formulated. The stages of the formation of a multi-level ballistic structure are considered in detail, from the standpoint of a systematic approach. A mathematical formulation of the problem and a hierarchy of performance indicators are proposed. For a meaningful description of the simulated system, a conceptual model for substantiating multi-level ballistic structures of a cluster orbital grouping of remote sensing of the Earth under the influence of the external environment has been developed. The model shows the interrelationships of the main elements of the substantiation of the ballistic structure of the cluster orbital grouping of remote sensing of the Earth and the sequence of formation of particular problems. The results of modeling both the ballistic structure of the cluster and the ballistic structure of the Earth remote sensing orbital grouping, which provides a set of tactical and technical, spatio-temporal and structurally stable consumer requirements, are presented. The effect of the application of the developed conceptual model will be the optimal strategy for the use of cluster orbital groupings of remote sensing of the Earth, which will provide the required value of its effectiveness under the influence of the external environment.

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Kondratiev, A. V. "A concept of optimization of structural and technological parameters of polymer composite rocket units considering the character of their production." Kosmìčna nauka ì tehnologìâ 26, no. 6 (2020): 5–22. http://dx.doi.org/10.15407/knit2020.06.005.

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We present a concept of optimization of structural and technological parameters of rocket and space technology units from polymer composite materials under heterogeneous loading and a project complex for their rational selection, taking into account the current level of production. The concept includes five interconnected components: design, production technologies, operation, ecology, and safety of industrial life. The analysis of possible criteria-based optimization estimates is carried out on the example of the technological component of the problem. Decompositions of the general task of parameters’ optimization were carried out into a number of types that correspond to the main types of structures of the considered class of technology: load-bearing compartments of launch vehicles and precision structures of spacecraft. An integrated approach to the optimal design of the bearing compartments of the head block of launch vehicles of various structural and power schemes is proposed. A distinctive feature of the approach is the possibility of multifactor optimization of the parameters for units of the class under consideration while providing regulated load-bearing capacity with simultaneous power and heat loading, taking into account technological, operational, economic, and environmental restrictions that correspond to the existing level of their production. A conceptual approach to the synthesis of rational parameters of composite frames of solar panels of various structural and power circuits is proposed, based on the integrated realization of well-known principles implemented by relevant units that are integrated by computer technology into a single optimization complex. An integrated approach has been synthesized to create precision space structures from polymer composite materials, which makes it possible to obtain rational thermo-dimensionally stable composite structures. An algorithm for determining the rational structure of a composite package has been developed and implemented, which provides a compromise combination for the absolute values of the coefficient of linear thermal expansion keeping maximum precision of the product in accordance with the proposed criteria. The results obtained made it possible to provide an increase by more than 20% in the mass efficiency of the composite aggregates of rocket and space technology produced at the leading enterprises of the industry.

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Cartmell, Matthew, Olga Ganilova, Eoin Lennon, and Gavin Shuttleworth. "Motorised momentum exchange space tethers: the dynamics of asymmetrical tethers, and some recent new applications." MATEC Web of Conferences 148 (2018): 01001. http://dx.doi.org/10.1051/matecconf/201814801001.

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This paper reports on a first attempt to model the dynamics of an asymmetrical motorised momentum exchange tether for spacecraft payload propulsion, and it also provides some interesting summary results for two novel applications for motorised momentum exchange tethers. The asymmetrical tether analysis is very important because it represents the problematic scenario when payload mass unbalance intrudes, due to unexpected payload loss or failure to retrieve. Mass symmetry is highly desirable both dynamically and logistically, but it is shown in this paper that there is still realistic potential for mission rescue should an asymmetry condition arise. Conceptual designs for tethered payload release from LEO and lunar tether delivery and retrieval are also presented as options for future development.

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LEVANDOVICH, Aleksandr V., Dmitry A. MOSIN, Viktor V. SINYAVSKIY, Aleksandr Ye TYUTYUKIN, and Igor A. UPTMINTSEV. "Conceptual design studies of an electrically propelled upper stage for deployment of a multi-plane orbital constellation of small spacecraft." Space engineering and technology, June 30, 2021, 97–108. http://dx.doi.org/10.33950/spacetech-2308-7625-2021-2-97-108.

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The paper presents results of conceptual design studies to determine configuration of an electrically propelled upper stage (EPUS) – a space transportation stage (a space tug) with main engines based on electric propulsion powered by solar arrays. It addresses the problem of deploying a multi-plane orbital constellation of small spacecraft (SSC) using an electrically propelled upper stage. It proposes to change the SSC operational orbital planes based on the effect of the difference in precession rates between the parking and the working orbits owing to the effect of eccentricity in the Earth gravitational field. Requirements have been defined for the EPUS electrical propulsion system that take into account the need to operate it to offset the aerodynamic drag while waiting in the parking orbit for the SSC operational orbital plane to turn. It demonstrates the feasibility of employing four EPUS that use Stationary Plasma Thruster-type electric propulsion as their main engines and gallium arsenide solar arrays for deployment in a 600 km orbit in four planes an orbital constellation of 24 small spacecraft with a mass of ~250 kg each using one launch of a medium capacity launch vehicle of Soyuz-2.1b type. Key words: Electrically propelled upper stage, electric propulsion, small spacecraft, orbital constellation.

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LEVANDOVICH, Aleksandr, Dmitry MOSIN, Aleksandr TYUTYUKIN, Igor URTMINTSEV, and Viktor SINYAVSKIY. "Conceptual design studies of an electrically propelled upper stage for deployment of a multi-plane orbital constellation of small spacecraft." Space engineering and technology, June 2021, 76–87. http://dx.doi.org/10.33950/spacetech-2308-7625-2021-2-76-87.

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The paper presents results of conceptual design studies to determine configuration of an electrically propelled upper (a space tug) with main engines arrays. It addresses the problem stage (EPUS) — a space transportation stage based on electric propulsion powered by solar of deploying a multi-plane orbital constellation of to in of small spacecraft (SSC) using an electrically propelled upper stage. It proposes change the SSC operational orbital planes based on the effect of precession rates between the parking and the working orbits eccentricity in the Earth gravitational field. Requirements owing have the difference to the effect been defined for the EPUS electrical propulsion system that take into account the need to operate it to offset the aerodynamic drag while waiting in the parking orbit for the SSC operational orbital plane to turn. It demonstrates the feasibility of employing four EPUS that use Stationary Plasma Thruster-type electric propulsion as their main engines and gallium arsenide solar arrays for deployment in a 600 km orbit in four planes an orbital constellation of 24 small spacecraft with a mass of ~250 kg each using one launch of a medium capacity launch vehicle of Soyuz-2.1b type.

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Kroll, Ehud, and Ido Farbman. "Casting innovative aerospace design case studies in the parameter analysis framework to uncover the design process of experts." Design Science 2 (2016). http://dx.doi.org/10.1017/dsj.2016.2.

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Traditional aerospace design methods offer quick and efficient ways to generate new designs, but such that often resemble previous ones. For truly innovative design, however, a different approach is needed. This paper suggests that a general conceptual design method called ‘parameter analysis’ (PA) may be used for teaching and practicing innovative aerospace design. To support this proposition, we investigate four diverse, innovative and unique case studies, all carried out by very experienced aerospace designers: the ‘dam busting’ bouncing bomb of World War II, the Gossamer Condor human-powered plane of the 1970s, the asymmetric Boomerang twin-engine plane of the 1990s and the SpaceShipOne suborbital spacecraft of the early 2000s. The paper elaborates on how the methodology of case-study research has been adapted and applied to provide the evidence supporting the research hypothesis, and presents the results of analyzing the case studies. It shows that the expert aerospace designers followed a thought process similar to PA, even if unknowingly, where the similarity was measured by counting the number of PA characteristics that could be shown to exist in the case studies. Advantages and limitations of the research methodology are also discussed.

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Gelain, Riccardo, Artur Elias De Morais Bertoldi, Adrien Hauw, and Patrick Hendrick. "3D Printing Techniques for Paraffin-Based Fuel Grains." Aerotecnica Missili & Spazio, September 1, 2022. http://dx.doi.org/10.1007/s42496-022-00126-5.

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AbstractHybrid rocket propulsion systems have proved to be a suitable option for some specific applications in the space transportation domain such as in launch vehicle upper stages, orbit transfer spacecrafts, decelerator engines for re-entry capsules, and small satellites launchers. Part of the renewed interest in hybrid rocket propulsion is due mainly to the safety aspects, cost reduction, and the use of paraffin-based fuel that impacts positively in terms of the solid fuel regression rate. However, paraffin solid fuel grains have poor structural characteristics and sometimes low performance due to the fuel internal ballistics behaviour. More recently, various studies have been carried out to overcome these drawbacks of paraffin-based fuels, such as the addition of energetic nano-sized metallic powder and 3D printing techniques. This study presents a review of the principal concepts of 3D printing processes and extrusion techniques that can be suitable for paraffin grains manufacturing and the conceptual design of a prototype for a 3D printer system under development at the Aero-Thermo-Mechanics Department of Université Libre de Bruxelles.

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PAVLOV, Aleksandr N., Dmitry A. PAVLOV, and Valentin N. VOROTYAGIN. "A method of using fuzzy hypergraphs to evaluate structural and technological survivability of attitude control systems for unmanned spacecraft." Space engineering and technology, September 30, 2020, 103–13. http://dx.doi.org/10.33950/spacetech-2308-7625-2020-3-103-113.

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Successful completion of a mission by an unmanned spacecraft (USC), both under nominal operational conditions both under examined contingencies and unexamined off-nominal situations is possible through designing survivability into the USC onboard system (OS). An analysis of current methods for evaluating USC OS survivability during their configuration management and reconfiguration under conditions of examined in-flight contingencies widely used in the design and development of the said USC has shown that these methods are not acceptable for evaluating the USC OS survivability in case of unexamined off-nominal situations in flight. This calls for development of conceptually novel methodological and procedural framework for evaluating structural survivability of USC OS configurations that take into account the level of participation of functional elements (FE) and OS subsystems in the USC control operations under various scenarios of the mission plan implementation. The paper proposes an original approach to evaluating the structural and technological survivability of the USC OS based on a fuzzy hypergraph formal representation of the operations to control the USC attitude, where the edges of the hypergraph connect the FE and OS subsystems that support the implementation of this or that specific control process. The paper also shows how one could use for the quantitative evaluation of the structural and technological survivability of a specific USC OS configuration the results of differentiation of a fuzzy hypergraph that could be visualized as a fuzzy hypergraph of technological independence of OS FE. Such an approach makes it possible to analyze the effects of FE on OS, identify the most critical elements, which have the lowest technological independence under mission plan implementation conditions, which could be used for providing a rationale for the required level of structural and functional redundancy of USC elements and subsystems introduced during various phases in its life cycle. Keywords: unmanned spacecraft, onboard systems survivability, fuzzy hypergraph derivative.

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Journal articles on the topic 'Spacecraft Conceptual Design'

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