МЕТОДИКА ОЦЕНКИ ДИАПАЗОНА ЭФФЕКТИВНОГО ПРИМЕНЕНИЯ УНИФИЦИРОВАННЫХ КОСМИЧЕСКИХ ПЛАТФОРМ

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Introduction. During a long period of space activity of JSC “ISS” a large number of spacecrafts (SC) of in- formation support and their component systems and in- struments have been developed. By doing so, a unified range of basic types of spacecraft (SCUR) was created, the modification of which allowed to create entire fami- lies of spacecraft for specific tasks and operating condi- tions quickly. At the same time, a technical policy was developed to introduce new principles and approaches in the creation of information space technologies, the main idea of which is expressed in the integration of various functions on one spacecraft; a significant increase in the life of the space- craft; increase of power capacity of the spacecraft; crea- tion of devices operating in a vacuum and spacecraft with a leaky instrument compartment; creation of unified space platforms (USP), unified on-board support systems and unified on-board instruments; reduction in the cost of development by replacing physical developmental models of spacecraft with software models and the use of equip- ment and software developed for various projects (inter- project unification), etc. However creation of a parametric series of unified space platforms is associated with the solution of the op- timization task of increasing the output effect of the spacecraft created on their basis in the presence of func- tional redundancy of the USP with simultaneous reduc- tion of the costs and terms of the spacecraft creation by minimizing the modifications of the USP. When creating a parametric series of SCUR and USP for spacecrafts providing the information support, this task was solved in a more empirical way, by summarizing the reserve of existing developments and assessing the feasibility and expedience of applying the existing reserve in the future (continuity of development). In this paper, a methodological basis for solving this problem is proposed. A formal description of a spacecraft project. The design of the spacecraft is carried out using models con- taining a number of parameters that are random variables with known or unknown distribution laws [1]. The sources of uncertainty are random factors of utili- zation, incompleteness of the initial data for the design, due to the error in forecasting the main technical, techno- logical, economic indicators, as well as the error in pre- dicting the conditions for the project due to the long dura- tion of its creation. For such models, the task of choosing the optimal de- sign of the spacecraft is transformed into the problem of choosing a solution under the conditions of uncertainty and is done by searching for such a project, which by tak- ing into account the uncertainty in a number of parame- ters delivers the extremum of the objective function whenever possible. As a result, the task of choosing the optimal project under uncertainty is solved in two stages: - at the first stage, a project that satisfies the condi- tions and constraints that determine the permissible range of SC existence, that is the permissible design of the spacecraft is developed; - at the second stage, the optimization of the parame- ters for the chosen criterion is carried out in a wide range of its existence, i. e, a quasi-optimal design of the space- craft is selected. At the same time, all restrictions on parameters should be satisfied with a high level of probability, which is a necessary and sufficient condition for the implementation of an acceptable project, that is ensuring its structural stability. The procedure of optimizing the project serves as guidelines and boils down to isolating the range of valid parameters, in which the efficiency index is close to opti- mal. This approach to design is called rational design [1]. Rational design clarifies and supplements the funda- mental principles of the system approach to the develop- ment of complex technical systems as follows: - in the synthesis of the system structure options, it is necessary to start from the uncertainty ranges of all the parameters and if these ranges overlap, then the alterna- tive is not considered; - completeness of mathematical models of the system and modeling errors should take into account uncertainty ranges in parameters; - when forming the optimization criterion, the quality indicators of the system are ranked according to the de- gree of their influence on the criterion, taking into account the reliability of their values; - comparison of different project variants is carried out under identical conditions of uncertainty. Thus, the main task of rational design is to provide conditions for the implementation of an acceptable project by ensuring that the critical parameters of spacecraft that are random variables are not exceeded by creating com- pensation mechanisms for these uncertainties ensuring a guaranteed existence of an acceptable project, i. e. struc- tural stability of the spacecraft project in the whole range of possible realizations of random parameters. When designing a spacecraft, the mechanism for par- rying uncertainties is reduced to the creation of central- ized reserves of spacecraft resources to parry uncertainties by its parameters and redistribution of these reserves as the project progresses. The choice of the nomenclature of critical parameters is carried out on the basis of an analysis of the most sig- nificant limitations that are associated with the problems the SC is to solve: - the solution of target tasks; - control of the spacecraft operation; - motion control of the spacecraft; - control of the angular position of the spacecraft; - maintenance of energy and heat balance; - ensuring compatibility of the spacecraft with a launch vehicle. According to the research made, the design parameters that determine the structural stability of a spacecraft (guaranteed satisfaction of constraints) include [1]: - the mass of the spacecraft and the mass of the work- ing body of the propulsion system (PS); - the volume of the spacecraft in the folded position, the volume of the instrument cluster and tanks of the PS, the area of solar batteries and a radiator, the dimensions of the antennas; - the eccentricity of the spacecraft mass center; - the moments of inertia of the spacecraft in the folded and working positions; - power consumption and heat release of the space- craft. At the same time, the mass, volume, power consump- tion of the spacecraft and its components are independent of the above nomenclature of parameters. Therefore, in order to implement an acceptable pro- ject, it is primarily planned to manage the budget for the mass and energy consumption of the spacecraft in the permissible range of change, as well as the formation of a layout scheme for the spacecraft that is resistant to changes in the parameters of the spacecraft. The limiting values of the mass and volume of the spacecraft are limited by the selected means of induction, and therefore the design of the spacecraft must be directed at their maximum use in order to increase the target effi- ciency. Project model of the spacecraft with USP. One of the effective mechanisms for implementing an acceptable project is the use of a modular-type layout scheme of a spacecraft consisting of a payload module (PM) and a unified space platform (USP) for which the mass and en- ergy budgets of the spacecraft are presented in the follow- ing form: МSC = МP + МUSP, WSC = WP + WUSP, (1) where МSC and WSC - are mass and power consumption of the spacecraft; МP and WP - are mass and energy con- sumption of the PM; МUSP and WUSP - are mass and power consumption of the USP. The budget of the spacecraft resources is formed on the basis of the maximum satisfaction of the payload re- quirements in the spacecraft resources (energy consump- tion mass, volume) in the form of a generalized payload mass MPg [1]: МPg = МP + KW · WP = MP·αP, αP = 1+ KW · WP/MP, (2) where αP - is the coefficient of partial costs of the SC resources to ensure the needs of the payload; KW - is the average coefficient of partial costs of the spacecraft mass for generating electricity and heat rejection, kg / W. In this case, the generalized mass of the payload MGP can be used to form the indicator of the spacecraft efficiency - the generalized coefficient of the partial costs of the spacecraft resources for the solution of the target task: KP = МPg / МSC = К0P · αP, (3) where К0P = МP / МSC - mass payload coefficient of a spacecraft. Costs for carrying out development work on the de- velopment of a spacecraft (CDW) according to the enlarged methodology are proportional to the costs for the manu- facture of a spacecraft (CM) [1]: СDW = КDW · СM. (4) The value of the coefficient КDW is determined by the novelty of the spacecraft being developed and its compo- nents, the volume of ground-based experimental testing of the spacecraft and its component parts, and is specified in the range 4-8. At the same time, for the spacecraft on the new USP КDW ≈ 8, and when using the borrowed USP КDW ≈ 4. Costs for the manufacture of a spacecraft, as a combi- nation of the costs of manufacturing its components and their integration into the spacecraft structure, depends on its target efficiency, reliability, mass, energy consump- tion, etc. Taking into account the fact that the mass of the spacecraft is limited by the power capabilities of the launch vehicle and is used to realize target tasks with a given efficiency and reliability, in design studies it is used as an equivalent to the cost of manufacturing the space- craft [1] СM = Сsi· МSC. (5) The value of the specific indicator Csi is determined on the basis of the statistical data processing on SC- analogues. Substituting equation (5) into equation (4), we obtain the functional dependence of the development work cost on the mass of the spacecraft. СDW = КDW · Сsi · МSC. (6) The obtained system of equations allows to formulate the criterion of the optimal SC project (objective function) of the scalar type, defined as the ratio of the efficiency index (МGP) to the cost indicator for the creation of the МPn = МSCn - МUSP = МSCn - (МSCb - МPb). Substituting the expression for МPn in the formula for calculating KPn we obtain: spacecraft CDW. МPg K K 0 ×a KPn = 1- МSCb × (1- K МSCn Pb + dW ) ; ESC = = P = P P . (7) W -W СDW КDW × Сsi КDW × Сsi dW = KW Pb Pn . (12) М Procedure for evaluating the effective use range of the USP. To create a modern spacecraft for various pur- poses in a sufficiently short time, it is advisable to use a unified space platforms (USP) [1-3]. The USP is intended for further installation and adaptation of the payload (P) on it and providing it with all the conditions for full-time operation and for the tasks set for SCb Substituting the equation for the calculation of KPn in inequality (10), we obtain a ratio for estimating the range of effective application of the USP (without its improve- ment): Pb 1- K КDWn the spacecraft. It should be noted that in practice the application area 1, 0 £ МSCb £ МSCn КDWb . (13) 1- KPb + dW of the USP without further development is very limited, which is due to the variability of payload (P) parameters (mass, power consumption, design), the use of different types of launch vehicles, operation orbits, etc. Therefore, there is often a need for modification and even substantial In the case of a connection between МPn and WP (coefficient αP), the expression for determining KPn and the mass ratios МSCb / МSCn will assume a different form: improvement of the USP for the specific characteristics of K = МPn ×a = é - МSCb æ1- KPb öù ×a , (14) a spacecraft. To exclude the need to improve the USP, it is worked Pn М SCn ПН ê1 ë ç a МSCn è ÷ú P P øû on the limiting characteristics of the PM and the space- 1- KPb × КDWn 1- K 0 КDWn craft as a whole. In this case, the target efficiency of the М a К Pb К 1, 0 £ SCb £ P DWb = DWb . (15) Pb spacecraft (KP) is somewhat reduced due to a reduction of the resources for the PM because of the availability of surplus resources for the USP. М SCn 1- KPb aP 1- K 0 Let us consider the case of the USP application devel- oped for a basic spacecraft, for a new spacecraft with a smaller mass and energy consumption. МSCb = МPb + МUSP; МSCn = МPn + МUSP; МSCn ≤ МSCb; WPn ≤ WPb; (8) The use of the USP on the new spacecraft reduces the cost of the DW, which leads to an increase in its ESC crite- Approbation of the procedure for evaluating the effective use range of the USP. The verification of the developed procedure for evaluating the effective use range of the USP is carried out using the example of USP “Express-1000NT” for geostationary spacecraft. Evaluation of effective use range of the USP will be car- ried out at КDWn / КDWb = 4/8 = 0.5 for the two calculation options [11]: ria. However, if the mass of the new spacecraft is differ- ent from the mass of the basic spacecraft (to a smaller side), the mass of its payload decreases and, accordingly, - for dependent values of MP, WP, using K0P (by formula (15)); a = 1+ К WP М P W P its efficiency decreases, which reduces its ESC criterion. The range of effective application of the USP on the new spacecraft is determined by a relative dimensionless crite- rion (index “b” refers to the base spacecraft, and the index “n” to the new spacecraft). dE = ESCb = KPb × КDWn £ 1, 0 . (9) ESCn KPn × КDWb For further research, we transform inequality (9) to the following form: KPn ³ КDWn . (10) KPb КDWb We will carry out calculation of the coefficient KPn us- ing the constancy of the USP mass for the base and new spacecraft: - for independent values of MP, WP, using KP and at δW = 0 (by formula (13)). Calculated data are given in see table, graphical repre- sentations of the effective use ranges of USP “Ekpress- 1000NT” are shown in see figure. The solid line denotes the use of К0P, the dotted line indicates KPb. The points on the graph indicate the realized spacecraft on the basis of the USP: 1) “Express-AT1” (weight 1672 kg); 2) “KAZSAT-3” (weight 1704 kg); 3) “TELKOM-3” (weight 1725 kg); 4) “Yamal-300K” (weight 1847 kg); 5) “ LYBID” (weight 1903 kg); 6) “AMOS-5” (weight 1929 kg) [12-15]. Conclusion. The results of the approbation allow us to make the conclusion that the developed procedure makes it possible to estimate the effective use range of unified KPn = МPn + KW ×WPn ; (11) МSCn space platforms for communication satellites in the geo- stationary orbit correctly. Effective use range of the unified space platform “Express-1000HT” № п/п Characteristics Value 1 Platform type E-1000NT 2 Basic satellite weight, kg MSC 1950 3 The maximum mass of the payload (P) (RTR + antennas), kg MP 500 4 Maximum payload power consumption, W WP 5900 5 Coefficient of energy efficiency, kg / W KW 0.048 6 Generalized payload mass MPg 783.2 7 The coefficient of generalized payload KPb 0.402 8 Payload coefficient К0 Pb 0.256 9 Coefficient of partial mass costs for payload power supply αP 1.566 10 The minimum mass of a new spacecraft using К0P MSCn 1633 11 The minimum mass of a new spacecraft using KP и δW = 0 MSCn 1459 Effective use range of the unified space platform “Express-1000HT” Диапазон эффективного применения УКП «Экспресс-1000НТ»

Об авторах

В. Е. Чеботарев

АО «Информационные спутниковые системы» имени академика М. Ф. Решетнёва»

Российская Федерация, 662972, г. Железногорск Красноярского края, ул. Ленина, 52

И. И. Зимин

АО «Информационные спутниковые системы» имени академика М. Ф. Решетнёва»

Email: i.zimin@iss-reshetnev.ru
Российская Федерация, 662972, г. Железногорск Красноярского края, ул. Ленина, 52

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