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Space Factors Influence on the Size Stability of Small Spacecraft Structure Elements
Available online at www.sciencedirect.com
ScienceDirect Procedia Engineering 185 (2017) 105 – 109
6th Russian-German Conference on Electric Propulsion and Their Application
Space factors influence on the size stability of small spacecraft structure elements Yu.V. Skvortsova,*, S.V. Glushkova, Chernyakin S.A.a a
Samara University, 34, Moskovskoye Shosse, Samara, 443086, Russia
Abstract This paper examines «Aist-2D» small spacecraft. Thermoelasticity problem is solved in two stages by finite element method using ANSYS® software. The first stage is thermal analysis. Space factors, devices and onboard equipment heat liberation are taken into account. External heat currents intensity and equipment work modes change depending on the spacecraft position on the Earth orbit. The second stage is stress-strain analysis of the structure, which is caused by the temperature field influence. As a result, the onboard equipment movement against structure elements and the acting forces are defined. © 2017 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license © 2017TheAuthors.PublishedbyElsevier B.V. (http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-reviewunder responsibilityof the organizing committee of RGCEP–2016. Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application Keywords:small spacecraft; space factors; thermal analysis; structural analysis; finite element method; onboard equipment; size stability.
1. Introduction A small spacecraft is affected by the combined influence of many physical processes, which are causes by radiation fields of the Sun, the Earth and the spaces, as well as the residual Earth atmosphere. The Sun radiation (which is usually treated as a point source of radiation with its light rays that are considered parallel) and the Earth radiation which leaves the Earth surface (the Sun radiation, reflected from the Earth atmosphere and surface, and the planet’s self-radiation) are considered to be the most important for insuring the size stability of the small spacecraft structure. In this work finite element method is applied through ANSYS® software. It provides unique opportunities for the interlinked multiphysics problems solving, which may combine, for example, strength and thermal physics not only
* Corresponding author. Tel.: +7-846-267-45-20. E-mail address: [email protected]
1877-7058 © 2017 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application
doi:10.1016/j.proeng.2017.03.326
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Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
within one program, but also within one model. The interaction between different physics fields in this case can be achieved either through direct relation or through load transfer. In the second case, the interaction is realized by using the results of one analysis as the loads in the other analysis. In the thermoelectricity problem it is assumed that the thermal and structural analysis are performed successively. Interdependent thermoelectricity equations describe body deformation which occurs under mechanical and thermal influence, as well as the reverse effect – its thermal field change due to deformation. However, in most cases temperature deformations are insignificant and, therefore, they have no significant influence on temperature distribution. It will be one-way link in this case. If one-way load transfer occurs, the following steps should be taken to solve thermoelasticity problem: - thermal problem defining and solving; - returning to preprocessor and modifying the database. The finite elements types should be changed (from thermal to structural ones), additional material properties should be set (such as Young’s modulus of elasticity, Poisson`s ratio, linear expansion coefficients etc.) and structural boundary conditions need to be defined; - reading the temperatures from the Jobname.RTH thermal analysis results file (LDREAD command); - structural problem solving, that is thermal strain and stress calculation of the structure.
2. Brief description of the structure «Aist-2D» small spacecraft structure is designed to accommodate target-oriented, scientific and supporting devices inside and outside the spacecraft. This design is intended to create and support both the supporting systems and the specified conditions for the normal functioning of all onboard systems at all stages of operation. The small spacecraft structure in the form of a rectangular parallelepiped is non-sealed and it consists of a frame, six external panels and one central panel which carries visual observation range electrooptical equipment. The frame is designed as a box-shaped section load-bearing structure made of aluminum alloy, which consists of three horizontally oriented structural rings, interconnected by struts. Each panel is a sandwich plate with a lightweight honeycomb core made of metal foil and two thin shells made of aluminum alloy. The layers are joined together with film adhesive. In the places where the panels are attached to the frame and in the mounting seats for the onboard equipment bushings made of D16 material are used. To ensure heat removal from the places where onboard equipment is installed all the panels except from the central one have embedded internal heat transfer tubes. The external heat transfer tubes of smaller diameter are installed perpendicular to them.
3. Thermal analysis The goal of the thermal analysis is temperature distribution calculation in the structure under investigation and the other output values which are associated with it. This analysis is performed using the commercial program ANSYS® which is based on finite element method. Two types of heat transfer are considered - thermal conductivity and heat emission. The following element types are used to build the thermal model. To idealize honeycomb panels a 4-node thermal element of the SHELL131 multilayered shell is used. For the three-layered shell with the linear temperature variation through the thickness of each layer (KEYOPT(3) = 1) the following options are taken as the nodal degrees of freedom: TBOT - temperature of the bottom surface; TE2 - temperature of the coupling face of the first and the second layer; NE3 - temperature of the coupling surface of the second and the third layer; TTOP - upper surface temperature. A 2-node thermally-conductive bar element LINK33 is used for the modeling of the frame, heat transfer tubes and fasteners. Its each node has only one TEMP (temperature) degree of freedom. An 8-node thermal solid body
Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
element SOLID278 is used to model devices and on-board equipment. It also has only on TEMP (temperature) degree of freedom at the node. Since the heating jacket element has degrees of freedom that are different from TEMP, the heat transfer between different parts of the model is carried out here using the equations which connect individual degrees of freedom. For example, for the heat transfer tubes the equations of connection can be written as follows: – for the internal tubes: TE2n = TEMPn; TE3n = TEMPn; – for the external tubes: TTOPn = TEMPn, where n – heat transfer tube node number, which is the same as the number of the corresponding panel node. The so-called radiosity solver method is used for modeling complex heat transfer by means of radiation between different surfaces of the structure under analysis and the open space. This method is supported by all 2- and 3dimensional elements of the ANSYS® program, which have temperature as a node degree of freedom. It should be noted that in the radiosity solver method the energy balance equation [1] which was obtained based on the Stefan–Boltzmann law is solved in conjunction with the main problem of thermal conductivity for the gray diffusion radiation analysis between two or more surfaces. ANSYS® software allows us to assign several open or closed enclosures at the same time, that is the systems of the surfaces which radiate on each other. These enclosures are used for the view factors calculation. The so-called hemicube method is applied for these purposes [2, 3]. Each open enclosure should have its own space temperature or its own space node which radiates at the temperature of the surrounding area. Three open enclosures (two internal and one external enclosure) are defined for the small satellite unsealed structure which is being analyzed. The open space temperature is set as the space temperature for the outer enclosure, while for the inner enclosures (given that they are closed) space nodes, located in these enclosures, are defined. In addition to radiation, the heat flux densities are also set as surface loads. They are caused by the full Sun emission and the Earth infrared radiation. They are calculated for each outer surface of the small satellite structure depending on its position on the Earth orbit. Moreover, the heat generation power is set as a volume load depending on the heat-generating and onboard equipment work modes. The example of temperature distribution on the outer surfaces of the panels (TTOP degree of freedom) at one moment of the small satellite flight is shown in fig. 1. It should be noted that one color, denoting 0оС temperature, shows the devices and onboard equipment, modeled with solid elements, lacking TTOP degrees of freedom. 4. Structural analysis As noted above, we must first switch the finite elements types in order to move from thermal to structural analysis. For example, the SHELL131 thermal shell elements, which are used for the honeycomb panels, should be replaced by the equivalent SHELL181 multilayered shell structural elements. The LINK33 elements, which model the frame and the internal heat-transfer tubes, can be replaced by the BEAM188 structural beam elements. At the same time, in order to define the sections for the frame the corresponding offsets should be set, that is the longitudinal axis offset of the finite element from the axis, which passes through the centers of gravity of the beam sections.
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Fig. 1. Temperature distribution (in оС ) along the outer surfaces of the panels at one point in time
It should be noted that for the thermal model areas which will not be used in the structural analysis the zero type of elements should be set. These elements will be ignored during the solving process. Since the external heattransfer tubes stiffness is insignificant, they may not be considered in the structural analysis. Moreover, the equipment thermal models volume elements should be set as zero type, since for their modelling in the structural analysis the MASS21 concentrated mass structural elements are used, which are attached to the panels by means of the rigid beam element MPC 184 (KEYOPT(1) = 1). Prohibition for any displacements of the node, located in the center of mass of the visual observation range electrooptical equipment is treated as the structural boundary conditions here. Temperature distribution, which was obtained in the first step, is used as a load here. As previously mentioned, the reading of the nodal temperature values from any loading steps (or for any moments in time) of the thermal analysis is carried out using the LDREAD command, the load passes from one node to another. The example of the temperature translational movement field is shown in fig. 2. It corresponds to temperature distribution on fig. 1. As a matter of fact, taking into consideration the boundary conditions of the model, these are the displacements in relation to the center of mass of the visual observation range electrooptical equipment.
Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
Fig. 2. Temperature translational motion field (in meters) at one point in time
5. Conclusions The above method of thermoelasticity problem solving for the small satellite structure allows to define onboard equipment displacements relative to the structure elements and the acting forces. It gives a chance to assess the influence of space factors on the size stability of a small satellite element structure. Thus, the use of the proposed approach makes it possible to design the structural and technological procedures aiming at ensuring the specified size stability of the small satellite structure. References [1] Siegal, R. Thermal Radiation Heat Transfer [Text] / R. Siegal and J. R. Howell. – New York : Hemisphere Publishing Corporation. – 1992. – 1072 p. [2] Glass, M.W. Chaparral – A library package for solving large enclosure ra-diation heat transfer problems [Text] / M.W. Glass. – Sandia National La-boratories. Albuquerque, NM. – 1995. [3] Cohen, M.F. The Hemi-Cube: A Radiosity Solution for Complex Environments [Text] / M.F. Cohen, D.P. Greenberg // Computer Graphics. – 1985. – Vol. 19, No. 3. – P. 31-40.
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ScienceDirect Procedia Engineering 185 (2017) 105 – 109
6th Russian-German Conference on Electric Propulsion and Their Application
Space factors influence on the size stability of small spacecraft structure elements Yu.V. Skvortsova,*, S.V. Glushkova, Chernyakin S.A.a a
Samara University, 34, Moskovskoye Shosse, Samara, 443086, Russia
Abstract This paper examines «Aist-2D» small spacecraft. Thermoelasticity problem is solved in two stages by finite element method using ANSYS® software. The first stage is thermal analysis. Space factors, devices and onboard equipment heat liberation are taken into account. External heat currents intensity and equipment work modes change depending on the spacecraft position on the Earth orbit. The second stage is stress-strain analysis of the structure, which is caused by the temperature field influence. As a result, the onboard equipment movement against structure elements and the acting forces are defined. © 2017 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license © 2017TheAuthors.PublishedbyElsevier B.V. (http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-reviewunder responsibilityof the organizing committee of RGCEP–2016. Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application Keywords:small spacecraft; space factors; thermal analysis; structural analysis; finite element method; onboard equipment; size stability.
1. Introduction A small spacecraft is affected by the combined influence of many physical processes, which are causes by radiation fields of the Sun, the Earth and the spaces, as well as the residual Earth atmosphere. The Sun radiation (which is usually treated as a point source of radiation with its light rays that are considered parallel) and the Earth radiation which leaves the Earth surface (the Sun radiation, reflected from the Earth atmosphere and surface, and the planet’s self-radiation) are considered to be the most important for insuring the size stability of the small spacecraft structure. In this work finite element method is applied through ANSYS® software. It provides unique opportunities for the interlinked multiphysics problems solving, which may combine, for example, strength and thermal physics not only
* Corresponding author. Tel.: +7-846-267-45-20. E-mail address: [email protected]
1877-7058 © 2017 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application
doi:10.1016/j.proeng.2017.03.326
106
Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
within one program, but also within one model. The interaction between different physics fields in this case can be achieved either through direct relation or through load transfer. In the second case, the interaction is realized by using the results of one analysis as the loads in the other analysis. In the thermoelectricity problem it is assumed that the thermal and structural analysis are performed successively. Interdependent thermoelectricity equations describe body deformation which occurs under mechanical and thermal influence, as well as the reverse effect – its thermal field change due to deformation. However, in most cases temperature deformations are insignificant and, therefore, they have no significant influence on temperature distribution. It will be one-way link in this case. If one-way load transfer occurs, the following steps should be taken to solve thermoelasticity problem: - thermal problem defining and solving; - returning to preprocessor and modifying the database. The finite elements types should be changed (from thermal to structural ones), additional material properties should be set (such as Young’s modulus of elasticity, Poisson`s ratio, linear expansion coefficients etc.) and structural boundary conditions need to be defined; - reading the temperatures from the Jobname.RTH thermal analysis results file (LDREAD command); - structural problem solving, that is thermal strain and stress calculation of the structure.
2. Brief description of the structure «Aist-2D» small spacecraft structure is designed to accommodate target-oriented, scientific and supporting devices inside and outside the spacecraft. This design is intended to create and support both the supporting systems and the specified conditions for the normal functioning of all onboard systems at all stages of operation. The small spacecraft structure in the form of a rectangular parallelepiped is non-sealed and it consists of a frame, six external panels and one central panel which carries visual observation range electrooptical equipment. The frame is designed as a box-shaped section load-bearing structure made of aluminum alloy, which consists of three horizontally oriented structural rings, interconnected by struts. Each panel is a sandwich plate with a lightweight honeycomb core made of metal foil and two thin shells made of aluminum alloy. The layers are joined together with film adhesive. In the places where the panels are attached to the frame and in the mounting seats for the onboard equipment bushings made of D16 material are used. To ensure heat removal from the places where onboard equipment is installed all the panels except from the central one have embedded internal heat transfer tubes. The external heat transfer tubes of smaller diameter are installed perpendicular to them.
3. Thermal analysis The goal of the thermal analysis is temperature distribution calculation in the structure under investigation and the other output values which are associated with it. This analysis is performed using the commercial program ANSYS® which is based on finite element method. Two types of heat transfer are considered - thermal conductivity and heat emission. The following element types are used to build the thermal model. To idealize honeycomb panels a 4-node thermal element of the SHELL131 multilayered shell is used. For the three-layered shell with the linear temperature variation through the thickness of each layer (KEYOPT(3) = 1) the following options are taken as the nodal degrees of freedom: TBOT - temperature of the bottom surface; TE2 - temperature of the coupling face of the first and the second layer; NE3 - temperature of the coupling surface of the second and the third layer; TTOP - upper surface temperature. A 2-node thermally-conductive bar element LINK33 is used for the modeling of the frame, heat transfer tubes and fasteners. Its each node has only one TEMP (temperature) degree of freedom. An 8-node thermal solid body
Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
element SOLID278 is used to model devices and on-board equipment. It also has only on TEMP (temperature) degree of freedom at the node. Since the heating jacket element has degrees of freedom that are different from TEMP, the heat transfer between different parts of the model is carried out here using the equations which connect individual degrees of freedom. For example, for the heat transfer tubes the equations of connection can be written as follows: – for the internal tubes: TE2n = TEMPn; TE3n = TEMPn; – for the external tubes: TTOPn = TEMPn, where n – heat transfer tube node number, which is the same as the number of the corresponding panel node. The so-called radiosity solver method is used for modeling complex heat transfer by means of radiation between different surfaces of the structure under analysis and the open space. This method is supported by all 2- and 3dimensional elements of the ANSYS® program, which have temperature as a node degree of freedom. It should be noted that in the radiosity solver method the energy balance equation [1] which was obtained based on the Stefan–Boltzmann law is solved in conjunction with the main problem of thermal conductivity for the gray diffusion radiation analysis between two or more surfaces. ANSYS® software allows us to assign several open or closed enclosures at the same time, that is the systems of the surfaces which radiate on each other. These enclosures are used for the view factors calculation. The so-called hemicube method is applied for these purposes [2, 3]. Each open enclosure should have its own space temperature or its own space node which radiates at the temperature of the surrounding area. Three open enclosures (two internal and one external enclosure) are defined for the small satellite unsealed structure which is being analyzed. The open space temperature is set as the space temperature for the outer enclosure, while for the inner enclosures (given that they are closed) space nodes, located in these enclosures, are defined. In addition to radiation, the heat flux densities are also set as surface loads. They are caused by the full Sun emission and the Earth infrared radiation. They are calculated for each outer surface of the small satellite structure depending on its position on the Earth orbit. Moreover, the heat generation power is set as a volume load depending on the heat-generating and onboard equipment work modes. The example of temperature distribution on the outer surfaces of the panels (TTOP degree of freedom) at one moment of the small satellite flight is shown in fig. 1. It should be noted that one color, denoting 0оС temperature, shows the devices and onboard equipment, modeled with solid elements, lacking TTOP degrees of freedom. 4. Structural analysis As noted above, we must first switch the finite elements types in order to move from thermal to structural analysis. For example, the SHELL131 thermal shell elements, which are used for the honeycomb panels, should be replaced by the equivalent SHELL181 multilayered shell structural elements. The LINK33 elements, which model the frame and the internal heat-transfer tubes, can be replaced by the BEAM188 structural beam elements. At the same time, in order to define the sections for the frame the corresponding offsets should be set, that is the longitudinal axis offset of the finite element from the axis, which passes through the centers of gravity of the beam sections.
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Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
Fig. 1. Temperature distribution (in оС ) along the outer surfaces of the panels at one point in time
It should be noted that for the thermal model areas which will not be used in the structural analysis the zero type of elements should be set. These elements will be ignored during the solving process. Since the external heattransfer tubes stiffness is insignificant, they may not be considered in the structural analysis. Moreover, the equipment thermal models volume elements should be set as zero type, since for their modelling in the structural analysis the MASS21 concentrated mass structural elements are used, which are attached to the panels by means of the rigid beam element MPC 184 (KEYOPT(1) = 1). Prohibition for any displacements of the node, located in the center of mass of the visual observation range electrooptical equipment is treated as the structural boundary conditions here. Temperature distribution, which was obtained in the first step, is used as a load here. As previously mentioned, the reading of the nodal temperature values from any loading steps (or for any moments in time) of the thermal analysis is carried out using the LDREAD command, the load passes from one node to another. The example of the temperature translational movement field is shown in fig. 2. It corresponds to temperature distribution on fig. 1. As a matter of fact, taking into consideration the boundary conditions of the model, these are the displacements in relation to the center of mass of the visual observation range electrooptical equipment.
Yu.V. Skvortsov et al. / Procedia Engineering 185 (2017) 105 – 109
Fig. 2. Temperature translational motion field (in meters) at one point in time
5. Conclusions The above method of thermoelasticity problem solving for the small satellite structure allows to define onboard equipment displacements relative to the structure elements and the acting forces. It gives a chance to assess the influence of space factors on the size stability of a small satellite element structure. Thus, the use of the proposed approach makes it possible to design the structural and technological procedures aiming at ensuring the specified size stability of the small satellite structure. References [1] Siegal, R. Thermal Radiation Heat Transfer [Text] / R. Siegal and J. R. Howell. – New York : Hemisphere Publishing Corporation. – 1992. – 1072 p. [2] Glass, M.W. Chaparral – A library package for solving large enclosure ra-diation heat transfer problems [Text] / M.W. Glass. – Sandia National La-boratories. Albuquerque, NM. – 1995. [3] Cohen, M.F. The Hemi-Cube: A Radiosity Solution for Complex Environments [Text] / M.F. Cohen, D.P. Greenberg // Computer Graphics. – 1985. – Vol. 19, No. 3. – P. 31-40.
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