جستجوی مقالات مرتبط با کلیدواژه « طراحي مفهومي » در نشریات گروه « مکانیک »

In general, according to the Russian Gost standard, the space propulsion engine system includes injection subsystems, thrust chamber, heat transfer and cooling system, launch system, feeding system, thrust vector control system which due to overlap and penetration The manifolds between the engine reservoirs and the fuel system are usually designed as a pair. In the conducted investigations, in addition to knowing the elements of the circuit and how to arrange the circuit, it also leads to knowing the requirements, goals and main requirements of the system design. Then, based on the studies, the circuit design of the feeding system and the selection of its elements have been done.Considering the cryogenic of the propellants in this project, design considerations for routes and tanks will be more important, and finally, sensitivityanalysis and conceptualdesign considerations have been presented.In this article, the types of feeding systems, Ahead problems and challenges are discussed and finally model of this subsystem will be presented based on the TRL

One of the satellite telecommunication services is the S&F service. In this article, we analyze the service characteristics and explain the impact of them on the design parameters. One of the challenges of telecommunication payload design is the mapping of service parameters to the technical requirements of payload components. First, analytical relationships between service parameters( the number of users, traffic and delay) and technical requirements( network capacity and on-board memory) are presented. Then the protocol stack and their parameters are selected with the aim of optimizing the power and network efficiency considering the link conditions. By identifying the service specifications impact on the design parameters and the design process systematic, payload requirements in order to develop network is easily extracted. One of the effective factors in the payload block diagram design is its reliability. The design of redundancy is proposed by considering the components reliability and the its function with the aim of achieving reliability and reducing fault propagation.

Achieving to new technologies with high reliability, along with reducing the cost and time of the design cycle, is one of the most important challenges in complex systems. In this paper, reliability based design of a space system is discussed in the conceptual design phase. Normally, there are eight steps in the design for reliability. The first step, planning, and the next seven steps that applied for the liquid propellant engine with electro-pump technology included: determination of failure modes; reliability modeling; reliability allocation; propagation of uncertainty; Implementation of the chosen method in reliability analysis; reliability prediction and reliability evaluation. Therefore, in this research has been performed to achieved method and implementation steps of reliability design in the conceptual design phase of a space system.

This study presents the conceptual design of an unmanned multirotor with a novel configuration whose main application is disaster management. This multirotor can carry 800 kg of cargo for a range of 400 km, and its forward flight speed is 100 km/h. A multirotor with the specified performance parameters has not yet been developed. Due to the stipulated performance, the electric propulsion system is inapplicable, and a hybrid propulsion system is considered for this multirotor. The hybrid propulsion system without energy conversion was selected for the proposed design after further investigating alternative hybridization approaches. Multirotor subsystems are conceptually designed based on take-off weight, i.e., take-off weight is first estimated, and then multirotor subsystems are designed based on it. Finally, the overall subsystems’ weight is compared to the predicted take-off weight. If there is a significant discrepancy, the procedure is repeated until the two values converge. This paper presents the last design loop, and the specifications of the electric propulsion system, fuel propulsion system, power transmission system, body structure, and skid are calculated.

This article presents a novel and comprehensive approach to conceptual design of expander cycle liquid engines. One advantage of the computational code developed is the possibility of considering the working regime without cavitation of pumps by maximizing specific impulse of the engine system. Based on the proposed algorithm, the steady state regime of the mathematical modeling engine becomes static. This is important to increase reliability and to reduce costs during conceptual design phase, as well as optimization. One of the factors affecting the reliability of a system is its simplicity, which eliminates complexity. Here, the generator has been eliminated. In addition to the dimensional and functional estimations of the engine system, the proposed algorithm examines the performance of the propulsion system (engine specific impulse, turbine efficiency, and pumps) at various work stations. The computational algorithm of the static mathematical model was developed in Microsoft Visual Studio language.

ABSTRACT: From the perspective of synthesis, system design is comprised of conceptual, preliminary,and detailed design. Conceptual design is the first and most important phase of the system design that could influence the quality of the product. In general, conceptual design is iterative processes that tries to satisfy identified requirements. The outcome of the conceptual design phase is one or more concepts which does not necessarily accompany any detail. If the selected configuration could not satisfy the requirements, the designer should make changes in the decisions or may change the selected configuration completely. This activity makes the conceptual design process longer and increases the cost of the design process,as well. In addition, in the last decade, product design experienced fundamental changes in its concept from focusing on performance, function, and durability to sustainable design criteria such as being environmentally friendly, considering global warming, reducing energy consumption, and conducting end-of-product life cycle management such as reusing, recycling and remanufacturing. These conduct the designer to try to improve conceptual design process with the adoption of new methodologies. This article intends to use axiomatic design methodology for an innovative design, reduce repetition and satisfies most of the requirements in the conceptual design process.

This paper presents a methodology for conceptual design of long endurance stratospheric airships and establish a baseline of specifications for a conventional configuration stratosphere airship, according to given performance and operational requirements. The methodology is validated by other design concepts, which previously developed for similar missions. The shape optimization of airship was introduced into the design process, and several optimum objectives can be selected including minimum drag, minimum surface area and minimum weight. Also, a multi-objective function was used to take account of various factors which influence airship subsystems- e.g. aerodynamics, structures, energy and weight- to determine the optimal shape of airship. An algorithm for generating the shape is developed and appropriate mathematical models for subsystems are constructed. Simulation results show the optimized shape gives an improvement in the multi-objective function compared with a reference shape. The baseline specifications of stratosphere airships designed for various shapes by this methodology are presented.

The design of combustor has long been the most challenging portion in the design process of a gas turbine. This paper focused on the conceptual design methodology for aircraft combustors. The necessity of this work arose from an urgent need for a comprehensive model that can quickly provide data in the initial phases (conceptual design and preliminary design) of the design process. The proposed methodology integrated the performance and the design of combustors. To accomplish this, a computer code has been developed based on the design procedures. The design model could provide the combustor geometry and the combustor performance. Based on the available inputs data in the initial phases of the design process, a chemical reactor network (CRN) approach is selected to model the combustion with a detailed chemistry. In this way, three different chemical mechanisms are studied for Jet-A aviation fuel. Furthermore, the droplet evaporation for liquid fuel and the non-uniformity in the fuel-air mixture are modelled. The results of a developed design tool are compared with data of an annular engine’s combustor. The results have good agreement with the actual geometry and outputs of engine test rig emissions.

The major purpose of this paper is to illustarte of statistical design accuracy using trajectory simulation for launch vehicles design in conceptual design phase and also sensitivity analysis of velocity relative to effective external forces. Considering the advantages of statistical design to prevent the time and cost losses, system specification of sample launch vehicle calculated based on statistical data of the studied population. Then, by solving the equations of motion, design parameters are calculated in such a way that difference of the final velocity of trajectory simulation and needed orbital speed is less than 1 percent. Studied launch vehicles are two-stage liquid propellant vehicles, with Portability 2.5-3.5 tons mass to the low earth orbit. To validate, curves of speed, altitude and angle of path of launch vehicle designed with statistical method, compared with curves of Tsiklon launch vehicle, therefore correct operation the mission and accuracy of the statistical design algorithm is proved. By comparing ideal speed and speed of simulation, speed changes of any effective force obtained. Eventually speed loss factor at each stage and sensitive percent of each stage speed relative to the force, for both launch vehicles, statistical design and tsiklon, is analyzed.

This paper proposes a supplementary method for conceptual design of satellite electrical power subsystem(EPS). Each of represented methods for satellite electricalpower subsystemconceptual design in different references have some advantages and also disadvantages, besides in each of the methods a determined part of this subsystem has been in focused. In this research, first advantages and disadvantages of existing approaches for the conceptual design of electrical power subsystemwere reviewed, continued with combining of previous methods, improved relationships and using some of the simulation methods plus the using of statistical databases, a complementary method with more ascendency and less disadvantages in comparison with other approaches was presented. Finally, using a data from a specific satellite and the results of the statistical design, the complementary method has been validated.

In the current article, an efficient hull form is presented to achieve 80 knot speed and its initial conceptual design is carried out. Presented hull form is able to gain 80 knot in sea state 3 and capable to satisfy the requirements related to payload and operational range. In this regards, a hard chine, deep V, warped and two stepped hull form is taken from Southampton’s series and by scaling the model to a 16 meters hull, various steps of conceptual design phase are evaluated. First, calm water resistance is analyzed and consequently, by using existing experimental data, sea-keeping in irregular waves is studied. Longitudinal stability is another issue which is surveyed in this article and no porpoising behavior is observed. Moreover, transverse stability of presented hull form is checked based on Blount criteria. Engine selection and weight distribution are final issues which are presented. Finally, by performing various steps of conceptual design phase, it is obviously clear that hard chine, deep V, warp and stepped hull form can satisfy existing requirement.

In this paper the problem of multidisciplinary and multi-objective conceptual design optimization of an air launched projectile (ALP) is investigated. The proposed task is performed using a three degree of freedom (DOF) flight dynamics simulation model and taking into account allthe constraintsinvolved in the optimization process. To maximize the payload weight as well as the range of ALP، the vehicle weight and balance، aerodynamics and stability disciplines are selected in the optimization process using a model with moderate levels of fidelity for each subject. The ALP design optimization problem contains 14 design variables and 2 target functions that include the payload weight and the vehicle range. Finally، a performance based comparison of results between the optimized ALP and its non-optimum initial configuration has been made. In addition، a Monte Carlo analysis is performed over the optimal ALP design to see the effects of launching uncertainties in meeting the mission requirements.

The purpose of this article is the conceptual design of a two-stage crew launch vehicle with side boosters. At first, in order to achieve a suitable design point, the statistical design technique is used, and then statistical design process is validated by using two degree of freedom simulation and doing energetic-mass calculations. Then, multidisciplinary design optimization approach is applied for initial conceptual design optimization. The preferred structure for multidisciplinary design optimization is all-at-once and Genetic Algorithm (GA) is used as the optimizer algorithm. In order to achieve more accuracy, Simplex method is employed using GA’s results as a combined algorithm. Having performed the optimization process, a mass decrease of about 4 tons from missile gross weight was attained with respect to normal simulation results, and as already known, the decrease in gross mass undeniably leads to a consequent decrease in the cost of producing and launching missiles.

Designing space propulsion systems as one of the important subsystems of the spacecrafts and upper stage space launch systems needs to bypass different and complicated steps. In this article the comprehensive process of designing liquid fuel low-thrust space propulsion systems was illustrated. In the presented pattern, first of all according to the requirements and mission constraints, the main characteristics of the system were determined and then other characteristics were extracted. Finally, for the evaluation of the presented pattern, a low-thrust space propulsion system was designed based on a special mission and the results were compared with a real model. Comparison between the designed space propulsion system and the real one showed an appropriate accuracy of the presented pattern.

In this article the results of research to achieve a comprehensive code of remote sensing satellite conceptual design is presented. In compiling the code with considering the design philosophy of "better, faster, cheaper" has been attempted in caddition to the use of new technologies and the experiences of experts, local constraints such as national launchers limitation also be considered.The main advantage of the proposed code than conventional models, is using accurate simulation methods and newer models in the process of conceptual design of subsystems. view of the practical experience of the past to choose the best design starting point, achieve an operational plan to reduce the risk of costly changes, next steps of design has been achieved.

T‌h‌e p‌r‌i‌n‌c‌i‌p‌l‌e g‌o‌a‌l o‌f t‌h‌i‌s p‌a‌p‌e‌r i‌s t‌o p‌r‌e‌s‌e‌n‌t a m‌e‌t‌h‌o‌d‌o‌l‌o‌g‌y f‌o‌r a r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y b‌a‌s‌e‌d l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e c‌o‌n‌c‌e‌p‌t‌u‌a‌l d‌e‌s‌i‌g‌n (R‌L‌V‌C‌D). I‌n t‌h‌i‌s d‌e‌s‌i‌g‌n m‌e‌t‌h‌o‌d‌o‌l‌o‌g‌y, l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y i‌s d‌e‌t‌e‌r‌m‌i‌n‌e‌d b‌a‌s‌e‌d o‌n c‌u‌s‌t‌o‌m‌e‌r n‌e‌e‌d‌s. T‌h‌e‌n, a m‌u‌l‌t‌i‌p‌l‌e w‌e‌i‌g‌h‌t‌e‌d c‌r‌i‌t‌e‌r‌i‌o‌n m‌e‌t‌h‌o‌d, c‌a‌l‌l‌e‌d t‌h‌e A‌n‌a‌l‌y‌t‌i‌c H‌i‌e‌r‌a‌r‌c‌h‌y P‌r‌o‌c‌e‌s‌s (A‌H‌P), i‌s e‌n‌a‌b‌l‌e‌d t‌o a‌s‌s‌e‌s‌s t‌h‌e r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y a‌l‌l‌o‌c‌a‌t‌i‌o‌n o‌f a l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e s‌u‌b‌s‌y‌s‌t‌e‌m, b‌a‌s‌e‌d o‌n u‌t‌i‌l‌i‌z‌e‌d t‌e‌c‌h‌n‌o‌l‌o‌g‌y, c‌o‌s‌t, c‌o‌m‌p‌l‌e‌x‌i‌t‌y a‌n‌d o‌p‌e‌r‌a‌t‌i‌o‌n t‌i‌m‌e. T‌h‌e‌s‌e c‌r‌i‌t‌e‌r‌i‌a a‌r‌e s‌e‌l‌e‌c‌t‌e‌d t‌o d‌e‌c‌i‌d‌e a‌b‌o‌u‌t t‌h‌e o‌p‌t‌i‌m‌u‌m r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y o‌f e‌a‌c‌h i‌n‌d‌i‌v‌i‌d‌u‌a‌l s‌u‌b‌s‌y‌s‌t‌e‌m, s‌u‌c‌h t‌h‌a‌t t‌h‌e r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y g‌o‌a‌l o‌f t‌h‌e l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e c‌a‌n b‌e a‌c‌h‌i‌e‌v‌e‌d. T‌h‌i‌s i‌s a l‌o‌g‌i‌c‌a‌l a‌p‌p‌r‌o‌a‌c‌h, b‌e‌c‌a‌u‌s‌e s‌y‌s‌t‌e‌m r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y i‌s t‌h‌e p‌r‌o‌d‌u‌c‌t o‌f s‌u‌b‌s‌y‌s‌t‌e‌m r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌i‌e‌s, a‌n‌d, t‌h‌e‌r‌e‌f‌o‌r‌e, i‌f t‌h‌e‌y a‌r‌e o‌p‌t‌i‌m‌i‌z‌e‌d a‌n‌d c‌o‌m‌p‌a‌t‌i‌b‌l‌e, w‌i‌t‌h r‌e‌s‌p‌e‌c‌t t‌o t‌h‌e m‌e‌n‌t‌i‌o‌n‌e‌d c‌r‌i‌t‌e‌r‌i‌a, s‌y‌s‌t‌e‌m r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y w‌i‌l‌l a‌l‌s‌o b‌e o‌p‌t‌i‌m‌i‌z‌e‌d a‌n‌d c‌o‌m‌p‌a‌t‌i‌b‌l‌e. T‌h‌e a‌f‌o‌r‌e‌m‌e‌n‌t‌i‌o‌n‌e‌d m‌e‌t‌h‌o‌d u‌t‌i‌l‌i‌z‌e‌s t‌h‌e l‌i‌n‌k b‌e‌t‌w‌e‌e‌n t‌h‌e M‌A‌T‌L‌A‌B c‌o‌d‌e (e‌v‌a‌l‌u‌a‌t‌i‌o‌n o‌f t‌h‌e i‌n‌c‌o‌n‌s‌i‌s‌t‌e‌n‌c‌y i‌n‌d‌e‌x o‌f a m‌a‌t‌r‌i‌x b‌a‌s‌e‌d o‌n E‌i‌g‌e‌n v‌e‌c‌t‌o‌r) a‌n‌d t‌h‌e E‌X‌C‌E‌L s‌h‌e‌e‌t (c‌r‌e‌a‌t‌i‌o‌n o‌f t‌h‌e c‌o‌m‌p‌a‌r‌e‌d m‌a‌t‌r‌i‌x o‌f o‌b‌j‌e‌c‌t‌i‌v‌e f‌u‌n‌c‌t‌i‌o‌n, c‌o‌m‌p‌a‌r‌e‌d m‌a‌t‌r‌i‌x o‌f c‌r‌i‌t‌e‌r‌i‌o‌n a‌n‌d c‌o‌m‌p‌a‌r‌e‌d m‌a‌t‌r‌i‌x o‌f a‌l‌t‌e‌r‌n‌a‌t‌i‌v‌e). I‌n t‌h‌e R‌L‌V‌C‌D m‌e‌t‌h‌o‌d‌o‌l‌o‌g‌y, t‌h‌e M‌o‌n‌t‌e C‌a‌r‌l‌o m‌e‌t‌h‌o‌d i‌s a‌p‌p‌l‌i‌e‌d t‌o l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y a‌n‌a‌l‌y‌s‌i‌s a‌n‌d w‌i‌l‌l a‌l‌s‌o v‌a‌l‌i‌d‌a‌t‌e t‌h‌e l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e s‌u‌b‌s‌y‌s‌t‌e‌m r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y a‌l‌l‌o‌c‌a‌t‌i‌o‌n. S‌o, i‌n t‌h‌e d‌e‌v‌e‌l‌o‌p‌e‌d a‌l‌g‌o‌r‌i‌t‌h‌m, m‌a‌s‌s-e‌n‌e‌r‌g‌y c‌o‌e‌f‌f‌i‌c‌i‌e‌n‌t‌s a‌n‌d t‌e‌c‌h‌n‌o‌l‌o‌g‌y s‌e‌l‌e‌c‌t‌i‌o‌n a‌r‌e a‌f‌f‌e‌c‌t‌e‌d b‌y s‌u‌b‌s‌y‌s‌t‌e‌m a‌l‌l‌o‌c‌a‌t‌e‌d r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y u‌s‌i‌n‌g a s‌t‌a‌t‌i‌s‌t‌i‌c‌a‌l p‌r‌o‌c‌e‌s‌s. I‌n t‌h‌i‌s p‌a‌p‌e‌r, o‌n‌l‌y t‌h‌e e‌f‌f‌e‌c‌t o‌f e‌n‌g‌i‌n‌e m‌o‌d‌i‌f‌i‌c‌a‌t‌i‌o‌n i‌n t‌h‌e d‌e‌s‌i‌g‌n p‌r‌o‌c‌e‌s‌s o‌n l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e b‌a‌l‌l‌i‌s‌t‌i‌c p‌a‌r‌a‌m‌e‌t‌e‌r‌s w‌i‌l‌l b‌e c‌o‌n‌s‌i‌d‌e‌r‌e‌d, d‌u‌e t‌o s‌e‌v‌e‌r‌a‌l p‌r‌o‌b‌l‌e‌m‌s, s‌u‌c‌h a‌s l‌a‌c‌k o‌f d‌a‌t‌a a‌b‌o‌u‌t o‌t‌h‌e‌r s‌u‌b‌s‌y‌s‌t‌e‌m‌s. T‌h‌e‌r‌e a‌r‌e s‌o‌m‌e s‌u‌b-c‌y‌c‌l‌e‌s i‌n t‌h‌e r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y b‌a‌s‌e‌d c‌o‌n‌c‌e‌p‌t‌u‌a‌l d‌e‌s‌i‌g‌n o‌f t‌h‌e l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e, i‌n w‌h‌i‌c‌h t‌h‌e a‌l‌g‌o‌r‌i‌t‌h‌m i‌t‌e‌r‌a‌t‌e‌s t‌o m‌e‌e‌t t‌h‌e d‌e‌s‌i‌g‌n r‌e‌q‌u‌i‌r‌e‌m‌e‌n‌t‌s. C‌o‌n‌s‌e‌q‌u‌e‌n‌t‌l‌y, t‌h‌e R‌L‌V‌C‌D m‌e‌t‌h‌o‌d‌o‌l‌o‌g‌y i‌s a‌p‌p‌l‌i‌e‌d t‌o a‌n e‌x‌i‌s‌t‌i‌n‌g l‌a‌u‌n‌c‌h v‌e‌h‌i‌c‌l‌e w‌i‌t‌h d‌i‌f‌f‌e‌r‌e‌n‌t r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y. T‌h‌e o‌b‌t‌a‌i‌n‌e‌d r‌e‌s‌u‌l‌t‌s s‌h‌o‌w t‌h‌a‌t t‌h‌e s‌u‌g‌g‌e‌s‌t‌e‌d m‌e‌t‌h‌o‌d‌o‌l‌o‌g‌y i‌s a‌n e‌f‌f‌i‌c‌i‌e‌n‌t m‌e‌t‌h‌o‌d f‌o‌r r‌e‌l‌i‌a‌b‌i‌l‌i‌t‌y b‌a‌s‌e‌d d‌e‌s‌i‌g‌n i‌n t‌h‌e c‌o‌n‌c‌e‌p‌t‌u‌a‌l d‌e‌s‌i‌g‌n p‌h‌a‌s‌e.

To design a manned spacecraft carrying one to two crews to the low Earth orbits, design phases should be completed in various levels. It also needs to gather manned spacecrafts technical data which is developed in the same category. In the system design algorithm presented in this paper, the conceptual design sequences of a manned spacecraft named Dousti is accomplished systematically. First of all, in accordance with a target group of manned spacecrafts’ mission, Dousti’s mission profile is defined and system level requirements are recognized. User’s requirements are also considered in the mission profile and subsequently in system level requirements. General characteristics of Dousti spacecraft as well as its mass and dimensional features are derived in the next step. Statistics and parametric models are systematically applied in design sequence. Then, final characteristics of the spacecraft’s main subsystems designed through engineering methods and applying parametric models are introduced. Afterwards, resulting characteristics of the spacecraft are traded off to reform and then validated by statistics and parametric models to present the final plan.

In this study, a design cycle of strap on booster is presented. This procedure, based on conceptual design of solid rocket motor and statistical studies of strap on booster, is derived. The conceptual design, usages the collaborative design that enables the user interfaces to affection the design parameters. Afterwards, the software of solid rocket design and simulation of missile flight has been produced.

Understanding the role and significance of the different components of the trunk neuro-musculo-skeletal system is of great importance for development of proper preventive and curative measures for low back pain. The experimental investigations, using lumbar moment measurement devices, are considered the most useful sources of information for detail understanding of the biomechanics of trunk and its components. This paper describes the conceptual design of a new lumbar moment measurement system, based on minimum deflection, anthropometric characteristics, and measurement considerations. The system is capable to measure the 3-D trunk moments using uni-directional torque sensors within a proper configuration of links and joints. Some other features of the system include convergence of its joints, high stability, and ability to adapt to a wide range of anatomical postures is standing and sitting positions.