per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14492
Satellite Conceptual Design Multi-Objective Optimization Using Co Framework
Satellite Conceptual Design Multi-Objective Optimization Using Co Framework
A. Jafarsalehi
ajafarsalehi@mail.kntu.ac.ir
1
M. Mirshams
2
R. Emami
3
This paper focuses upon the development of an efficient method for conceptual design optimization of a satellite. There are many option for a satellite subsystems that could be choice, as acceptable solution to implement of a space system mission. Every option should be assessment based on the different criteria such as cost, mass, reliability and technology contraint (complexity). In this research, mass and technology constraints, which have a direct impact on the satellite life cycle cost, are considerd as system level objective function to obtain the system optimal solution during the coceptual design phase. The approach adopted in this paper is based on a distributed collaborative optimization (CO) framework. At system level, multiobjective optimization goal is to minimize the dry mass of the satellite and, simultaneously, minimize the system technology complexity which is subject to equality constraints. The use of equality constraints at the system level in CO to represent the disciplinary feasible regions, introduces numerical and computational difficulties as the discipline level optima are non-smooth and noisy functions of the system level optimization parameters.To address these difficulties robust optimization algorithms such as genetic algorithms (GA) are used at the system level. The results show that the CO framework has the same level of accuracy as the conventional All-At-Once approaches.
This paper focuses upon the development of an efficient method for conceptual design optimization of a satellite. There are many option for a satellite subsystems that could be choice, as acceptable solution to implement of a space system mission. Every option should be assessment based on the different criteria such as cost, mass, reliability and technology contraint (complexity). In this research, mass and technology constraints, which have a direct impact on the satellite life cycle cost, are considerd as system level objective function to obtain the system optimal solution during the coceptual design phase. The approach adopted in this paper is based on a distributed collaborative optimization (CO) framework. At system level, multiobjective optimization goal is to minimize the dry mass of the satellite and, simultaneously, minimize the system technology complexity which is subject to equality constraints. The use of equality constraints at the system level in CO to represent the disciplinary feasible regions, introduces numerical and computational difficulties as the discipline level optima are non-smooth and noisy functions of the system level optimization parameters.To address these difficulties robust optimization algorithms such as genetic algorithms (GA) are used at the system level. The results show that the CO framework has the same level of accuracy as the conventional All-At-Once approaches.
http://jsst.ias.ir/article_14492_f338063358564329a5dffa787cb95985.pdf
Design optimization
multiobjective
collaborative optimization
satellite conceptual design
Genetic Algorithms
Design optimization
multiobjective
collaborative optimization
satellite conceptual design
Genetic Algorithms
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14493
Effects of Some Parameters on Thermal Control of a LEO Satellite
Effects of Some Parameters on Thermal Control of a LEO Satellite
F. Farhani
f.farhani@irost.ir
1
A. Anvari
2
Satellite thermal control ensures safe operating temperature ranges for satellite components throughout the mission life. Effects of altitude, spin, and position of satellite radiator(s) on the thermal control of a small Low Earth Orbit (LEO) satellite have been studied. Results show that change in satellite altitude, in the range considered here, does not produce critical thermal conditions. However, satellite spin rate has a marked influence on the satellite temperatures. Also, comparison of results for the satellite configurations considered in this study suggests that a radiator at top provides better thermal design conditions. Results also indicate the adequacy of the discussed considerations for use in the design of satellites of similar configurations, missions and orbital parameters.
Satellite thermal control ensures safe operating temperature ranges for satellite components throughout the mission life. Effects of altitude, spin, and position of satellite radiator(s) on the thermal control of a small Low Earth Orbit (LEO) satellite have been studied. Results show that change in satellite altitude, in the range considered here, does not produce critical thermal conditions. However, satellite spin rate has a marked influence on the satellite temperatures. Also, comparison of results for the satellite configurations considered in this study suggests that a radiator at top provides better thermal design conditions. Results also indicate the adequacy of the discussed considerations for use in the design of satellites of similar configurations, missions and orbital parameters.
http://jsst.ias.ir/article_14493_0ef7ce13a7aa7bcb4d86568f8ea40013.pdf
Altitude
beta angle
low earth orbit (LEO)
satellite radiator
spin rate
thermal control
Altitude
beta angle
low earth orbit (LEO)
satellite radiator
spin rate
thermal control
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14494
Reliability Determination of a Sounding Rocket Separation System Using its Reliability Block Diagram and FMEA
Reliability Determination of a Sounding Rocket Separation System Using its Reliability Block Diagram and FMEA
M. A. Farsi
farsi@ari.ac.ir
1
A. A. Eslami
2
R. Gorgin
3
Separation system is one of the most important systems in rockets. The influence of this system on mission success cannot be ignored. In this paper, reliability of a sounding rocket separation system is determined using block diagram and FMEA . This system is based on the flexible linear shape charge cross-section and a spring mechanism to accelerate separation. In this investigation, the reliability block diagram of the separation system including mechanical and electrical mechanisms is determined. By considering reliability of each component based on expert opinion and using separation system reliability block diagram, reliability of separation system is determined. Moreover, since spring mechanism is one of the most important parts of separation system, a complete FMEA analysis is conducted for this mechanism. According to this analysis, piston, cylinder, pin, and springs have the highest RPN number. Hence, these parts must have a high reliability. On the other hand, results are shown that bracket and bush have the lowest RPN number; therefore, it is not important for these parts to have a high reliability.
Separation system is one of the most important systems in rockets. The influence of this system on mission success cannot be ignored. In this paper, reliability of a sounding rocket separation system is determined using block diagram and FMEA . This system is based on the flexible linear shape charge cross-section and a spring mechanism to accelerate separation. In this investigation, the reliability block diagram of the separation system including mechanical and electrical mechanisms is determined. By considering reliability of each component based on expert opinion and using separation system reliability block diagram, reliability of separation system is determined. Moreover, since spring mechanism is one of the most important parts of separation system, a complete FMEA analysis is conducted for this mechanism. According to this analysis, piston, cylinder, pin, and springs have the highest RPN number. Hence, these parts must have a high reliability. On the other hand, results are shown that bracket and bush have the lowest RPN number; therefore, it is not important for these parts to have a high reliability.
http://jsst.ias.ir/article_14494_0b8a5fc1e29b012b140314f7069ff320.pdf
separation system
reliability block diagram
FMEA
separation system
reliability block diagram
FMEA
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14495
Laminar and Turbulent Aero Heating Predictions over Blunt Body in Hypersonic Flow
Laminar and Turbulent Aero Heating Predictions over Blunt Body in Hypersonic Flow
S. A. Hosseini
hosseini@ari.ac.ir
1
S. Noori
2
In the present work, an engineering method is developed to predict laminar and turbulent heating-rate solutions for blunt reentry spacecraft at hypersonic conditions. The calculation of aerodynamic heating around blunt bodies requires alternative solution of inviscid flow field around the hypersonic bodies. In this paper, the procedure is of an inverse nature, that is, a shock wave is assumed and calculations proceed along rays normal to the shock. The solution is iterated until the given body is computed. The inverse method is practical for the calculation of flow field between the shock wave and the body surface. Body calculation with the body analysis is contrasted and according to the entire differences between those; the shape of shock with the coefficient scales is implemented. The normal momentum equation is replaced with a Maslen’s second order pressure equation. This significantlysignificantly decreases machine computation time. The present method predicts laminar and turbulent heating-rates that compare favorably with other researches. Since the method is very high-speed, it can be used for preliminary design, or parametric study of aerodynamics vehicles and thermal protection of hypersonic flows.
In the present work, an engineering method is developed to predict laminar and turbulent heating-rate solutions for blunt reentry spacecraft at hypersonic conditions. The calculation of aerodynamic heating around blunt bodies requires alternative solution of inviscid flow field around the hypersonic bodies. In this paper, the procedure is of an inverse nature, that is, a shock wave is assumed and calculations proceed along rays normal to the shock. The solution is iterated until the given body is computed. The inverse method is practical for the calculation of flow field between the shock wave and the body surface. Body calculation with the body analysis is contrasted and according to the entire differences between those; the shape of shock with the coefficient scales is implemented. The normal momentum equation is replaced with a Maslen’s second order pressure equation. This significantlysignificantly decreases machine computation time. The present method predicts laminar and turbulent heating-rates that compare favorably with other researches. Since the method is very high-speed, it can be used for preliminary design, or parametric study of aerodynamics vehicles and thermal protection of hypersonic flows.
http://jsst.ias.ir/article_14495_bd662bce850666b5d78d6f12c0d4508a.pdf
Aerodynamic heating
blunt body
Hypersonic Flow
Flow Field
Turbulent Flow
Laminar Flow
Aerodynamic heating
blunt body
Hypersonic Flow
Flow Field
Turbulent Flow
Laminar Flow
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14496
Approximate Viscous Shock-Layer Analysis of Axisymmetric Bodies in Perfect Gas Hypersonic Flow
Approximate Viscous Shock-Layer Analysis of Axisymmetric Bodies in Perfect Gas Hypersonic Flow
S. Ghasemloo
1
S. Noori
s_noori@aut.ac.ir
2
In this paper, an approximate axisymmetric method is developed which can reliably calculate fully viscous hypersonic flow over blunt-nosed bodies. In this method, a Maslen’s second-order pressure expression is used instead of the normal momentum equation. The combination of Maslen’s second-order pressure expression and viscous shock layer equations is developed to accurately and efficiently compute hypersonic flow fields of perfect gas around blunt-body configurations. The results show that, this combination leads to more accurate solutions and less extensive computer run times in the preliminary design environment. Furthermore, the utility of Cebeci-Smith turbulence model is adequate for application to long slender bodies. The results of these computations are found to be in good agreement with available numerical and experimental data.
In this paper, an approximate axisymmetric method is developed which can reliably calculate fully viscous hypersonic flow over blunt-nosed bodies. In this method, a Maslen’s second-order pressure expression is used instead of the normal momentum equation. The combination of Maslen’s second-order pressure expression and viscous shock layer equations is developed to accurately and efficiently compute hypersonic flow fields of perfect gas around blunt-body configurations. The results show that, this combination leads to more accurate solutions and less extensive computer run times in the preliminary design environment. Furthermore, the utility of Cebeci-Smith turbulence model is adequate for application to long slender bodies. The results of these computations are found to be in good agreement with available numerical and experimental data.
http://jsst.ias.ir/article_14496_821a8c1de0024f58437d1849269342fe.pdf
Approximate
Aerodynamic heating
HypersonicFlow
Perfect Gas
Approximate
Aerodynamic heating
HypersonicFlow
Perfect Gas
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14497
Fuzzy Sliding Mode for Spacecraft Formation Control in Eccentric Orbits
Fuzzy Sliding Mode for Spacecraft Formation Control in Eccentric Orbits
A. Imani
1
M. Bahrami
mbahrami@aut.ac.ir
2
The problem of relative motion control for spacecraft formation flying in eccentric orbits is considered in this paper. Due to the presence of nonlinear dynamics and external disturbances, a robust fuzzy sliding mode controller is developed. The slopes of sliding surfaces of the conventional sliding mode controller are tuned according to error states using a fuzzy logic and reach the pre-defined slopes. The controller is designed based on the nonlinear model of relative motion and perturbation and atmospheric drag are considered as external disturbances. Using the Lyapunov second method, the stability of the closed-loop system is guaranteed. The performance of the presented controller in tracking the desired reference trajectory is compared to a sliding mode controller in which simulation results confirm the superior performance of the proposed controller.
The problem of relative motion control for spacecraft formation flying in eccentric orbits is considered in this paper. Due to the presence of nonlinear dynamics and external disturbances, a robust fuzzy sliding mode controller is developed. The slopes of sliding surfaces of the conventional sliding mode controller are tuned according to error states using a fuzzy logic and reach the pre-defined slopes. The controller is designed based on the nonlinear model of relative motion and perturbation and atmospheric drag are considered as external disturbances. Using the Lyapunov second method, the stability of the closed-loop system is guaranteed. The performance of the presented controller in tracking the desired reference trajectory is compared to a sliding mode controller in which simulation results confirm the superior performance of the proposed controller.
http://jsst.ias.ir/article_14497_ca6c6bd7d466630e0584aab06e1d7d29.pdf
Sliding mode
Control
spacecraft formation flying
eccentric orbits
Fuzzy
Sliding mode
Control
spacecraft formation flying
eccentric orbits
Fuzzy
per
پژوهشگاه هوافضا و انجمن هوافضای ایران،
شماره ثبت مجوز نشریه از وزارت فرهنگ و ارشاد اسلامی 82946 مورخ 1397/07/02
علوم و فناوری فضایی
2008-4560
2423-4516
2014-04-01
7
1
14498
FPGA Implementation of JPEG and JPEG2000-Based Dynamic Partial Reconfiguration on SOC for Remote Sensing Satellite On-Board Processing
FPGA Implementation of JPEG and JPEG2000-Based Dynamic Partial Reconfiguration on SOC for Remote Sensing Satellite On-Board Processing
A. Chekini
1
H. R. Naji
hamidnaji@ieee.org
2
This paper presents the design procedure and implementation results of a proposed hardware which performs different satellite Image compressions using FPGA Xilinx board. First, the method is described and then VHDL code is written and synthesized by ISE software of Xilinx Company. The results show that it is easy and useful to design, develop and implement the hardware image compressor using new techniques of programmable logic tools for space applications. In this paper the proposed hardware uses the proposed hardware, and it is put on board of satellite. Appropriate bit streams are produced by synthesis tools; therefore, we have two bit streams which can be configured at any moment of time according to the user request. When users intend the hardware is reconfigured and changed from JPEG to JPEG2000 or vice versa. The Proposed architecture has some advantages other than previous architectures such as high-speed and real-time processing, high flexibility, low cost, high security and low power consumption. This idea can be utilized in modern commercial hardware space board for data compressing due to using partial reconfiguration technique.
This paper presents the design procedure and implementation results of a proposed hardware which performs different satellite Image compressions using FPGA Xilinx board. First, the method is described and then VHDL code is written and synthesized by ISE software of Xilinx Company. The results show that it is easy and useful to design, develop and implement the hardware image compressor using new techniques of programmable logic tools for space applications. In this paper the proposed hardware uses the proposed hardware, and it is put on board of satellite. Appropriate bit streams are produced by synthesis tools; therefore, we have two bit streams which can be configured at any moment of time according to the user request. When users intend the hardware is reconfigured and changed from JPEG to JPEG2000 or vice versa. The Proposed architecture has some advantages other than previous architectures such as high-speed and real-time processing, high flexibility, low cost, high security and low power consumption. This idea can be utilized in modern commercial hardware space board for data compressing due to using partial reconfiguration technique.
http://jsst.ias.ir/article_14498_508fc06120c20add0e87440db9270539.pdf
compression
satellite
on
Board Processing
real
Time Processing
Reconfigurable
compression
satellite
on
Board Processing
real
Time Processing
Reconfigurable