UNIVERSITY OF SCIENCE AND TECHNOLOGY OF HANOI

UNDERGRADUATE SCHOOL

BACHELOR THESIS Study of IGOSAT Mechanical Conception By Student: Pham Van Phap Department of Space and Aeronautics

External supervisor: Assoc. Prof. Hubert Halloin Laboratoire AstroParticule et Cosmologie Internal supervisor : Dr. Nguyen Xuan Truong Coordinator-Department of Space and Aeronautics

Paris, August 2015

Content Content .....................................................................................................................................................ii Acknowledgement ................................................................................................................................... iv Documents ................................................................................................................................................v Applicable documents ..........................................................................................................................v List of acronyms ...................................................................................................................................v List of table.......................................................................................................................................... vi List of figure ....................................................................................................................................... vii Abstract ................................................................................................................................................. viii INTRODUCTION ................................................................................................................................... 1 I. MODELING IGOSAT ......................................................................................................................... 2 1. Assembling IGOSAT ...................................................................................................................... 3 1.1. UHF/VHF antenna ................................................................................................................... 3 1.2. Middle cube (second cube)....................................................................................................... 4 1.3. Bottom cube (first cube) ........................................................................................................... 5 2. Assembly procedure ........................................................................................................................ 7 3. IGOSat skeleton .............................................................................................................................. 8 II. MECHANICAL & ENVIRONMENTAL TESTING PROPOSAL ................................................... 9 1. Purposes .......................................................................................................................................... 9 2. Scenario ......................................................................................................................................... 10 2.1. Sinusoidal vibration test ......................................................................................................... 10 2.2. Random vibration test ............................................................................................................ 10 2.3. Shock test ............................................................................................................................... 10 2.4. Thermal vacuum test .............................................................................................................. 11 2.5. Thermal cycling test ............................................................................................................... 11 3. Test profile .................................................................................................................................... 11

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3.1. Acceleration (quasi-static) test ............................................................................................... 11 3.2. Resonance survey ................................................................................................................... 12 3.3. Sinusoidal vibration test ......................................................................................................... 12 3.4. Random vibration test ............................................................................................................ 12 3.5. Shock test ............................................................................................................................... 13 3.6. Thermal tests .......................................................................................................................... 13 III. SIMULATIONS RESULTS & DICUSSIONS ............................................................................... 15 1. Simulating GPS adapter ............................................................................................................ 16 2. Simulating Skeleton .................................................................................................................. 17 3. Simulating Bottom cube ............................................................................................................ 19 4. Simulating Top cube ................................................................................................................. 20 5. Simulation the whole satellite ................................................................................................... 23 IV. CONCLUSIONS ............................................................................................................................. 26 REFERENCE .................................................................................................................................... 27 Appendix A:

STRUCTURAL REQUIREMENTS ........................................................................ A-1

Fundamental requirements ............................................................................................................... A-1 Payloads requirements ...................................................................................................................... A-1 Deployment requirements ................................................................................................................ A-2 Adaptation for small modules .......................................................................................................... A-2 Mass budget...................................................................................................................................... A-2 Summary table.................................................................................................................................. A-3 Appendix B:

List of structure components .................................................................................... B-1

Appendix C:

Simulation Materials Properties ............................................................................... C-1

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Acknowledgement The internship opportunity I had with IGOSat project was a great chance for learning, developing professional skills. I, therefore, consider myself as a very lucky individual as I was a part of it, and for had a chance to meet so many wonderful people and professionals who came with me through this internship period. Bearing in mind previous, I would like to express my deepest gratitude and special thanks to Assoc. Prof. Hubert Halloin, who is the principal investigator of the project, and my supervisor. He is who in spite of being extraordinarily busy with his duties, took time out to hear, guide and keep me on the correct path and allowing me to carry out my project at their esteemed organization and extending during the training. I also express my special thanks to Marco Agnan, the project Manager for taking part in useful decision and giving necessary advices and guidance and arranged all facilities to make life easier. I choose this moment to acknowledge his contribution gratefully. With my best regards, I cannot accomplish my internship in both theoretically and practically without the careful, precious, and extremely valuable support of: Dr. Xuan Truong Nguyen, my internal supervisor from USTH, Mr. Alain Givaudan, Mr. Walter Bertoli from APC, Dr. Xuan Huy Le, Mr. Nam Duong Nguyen from VNSC, and all members of the IGOSat team.

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Documents Applicable documents Application no. [A01] [A02] [A03] [A04] [A05] [A06]

Document Name cds_rev13_final

Document Title Cubesat Design Specification Rev.13 General Environmental Verification Standard for GSFC GSFC-STD-7000 Flight Programs and Projects ELV, VEGA Qualification and Acceptance test of VG-SG-1-C-040-SYS Equipment, Issued 5, Revision 1 PC104_plus_spec PC/104 – Plus Specification Version 2.0 IGOsat_SM_31012014 SPECIFICATION MISSION IGOSAT STR-TV-02 IGOSat STM Integration Manual, Version 1.1

List of acronyms ADCS BRF CDS EM FEM FM g GEVS GSFC IGOSAT ISIS NASA Oct/min PCB P-POD/ISIS-POD PSD QM s STM TBD UHF/VHF

Attitude Determination and Control System Body Reference Frame Cubesat Design Specification Engineering Model Finite Element Method Flight Model Gravitational Acceleration General Environmental Verification Standard Goddard Space Flight Center Ionosphere and Gamma-ray Observations Satellite Innovative Solution In Space National Aeronautics and Space Administration Octave per minute Printed Circuit Board Poly/ISIS Picosatellite Orbital Deployer Power Spectrum Density Qualification Model Second Structure and Thermal Model To be determined Ultra High Frequency/Very High Frequency

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List of table Table 1: QB50 Sinusoidal vibration test characteristics Table 2: QB50 Random vibration test characteristics Table 3: QB50 Shock test characteristics Table 4: Thermal vacuum test characteristics Table 5: Thermal cycling test characteristics Table 6: Simulating GPS adapter of 1 PCB summary Table 7: Simulating GPS adapter of 2 PCBs summary Table 8: Simulating skeleton without carbon plates summary Table 9: Simulating skeleton and carbon plates summary Table 10: Simulating Bottom cube summary Table 11: Simulating Top cube summary Table 12: Simulating Simplified IGOSat Table A-1: Mass budget summary of sub-systems

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List of figure Figure 1: IGOSat CAD Models Figure 2: Connection between antenna and structure Figure 3: Organization in middle cube Figure 4: The OEM 615 GPS board and its adaption on PCB Figure 5: Bottom cube layout Figure 6: The OEM 615 GPS board and its modified adaption on PCB Figure 7: Modified rib Figure 8: Bottom cube plan B’s layout: side view (left) and isometric view (right) Figure 9: IGOSat Skeleton Figure 10: Acceleration deformation GPS adapter. Figure 11: Deformation (maximum) result of acceleration simulation of skeleton Figure 12: Deformation (maximum) result of acceleration simulation of skeleton with carbon fiber plates Figure 13: Deformation (maximum) result of acceleration simulation of bottom cube Figure 14: Deformation (maximum) result of random vibration simulation of bottom cube Figure 15: Deformation (maximum) result of acceleration simulation of top cube Figure 16: Deformation (maximum) result of random vibration simulation of top cube Figure 17: IGOSat acceleration simulation on Z axis Figure 18: IGOSat simulated natural frequencies Figure 19: Random vibration's Power spectrum densities on three axes Figure 20: Random vibration's Power spectrum densities on Z axis

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Abstract Since the Cubesat has been propsed in 1999 by Professors Jordi Puig-Suari (California Polytechnic State University) and Bob Twiggs (Stanford University), there are hundreds of these nano-satellites were launched and are being developed. Beside of the goal to study space around the Earth with a little amount of money, the other essential expectation is for educational purposes that enable students’ ability to design, build, test, and operate in space spacecrafts. With this ambition, the IGOSat project, which is under the supervisor of University Paris Diderot, allows students to gain their practical experiences on several phases of satellite technologies. As worked with the Structure Subsystem, the roadmap of this thesis went from model, visualize, and simulate, with professional 3D CAD software, the mechanical structure of the satellite; take into account the various constraints of the different subsystems; to construct, mount and test the first prototype. Due to several standards and criteria that need to follow, the work was done with the collaboration with other subsystems to making decisions and modifications. At undergraduate level, this thesis can be considered as succeed in learning, practicing how to apply methods for advanced components and structure prototyping; perceiving the complexity of this field. On the other hand, the useful of this works for the Project is still being examined carfully. Keyword: IGOSat, 3U satellite, student satellite, structure subsystem, mechanical conception, CATIA.

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INTRODUCTION IGOSat, which stands for Ionospheric Gamma Ray Observation Satellite, is a CubeSat project being developed by Paris-Diderot University, with the support of Project JANUS and the Laboratoire d'Excellence UnivEarthS, and the participations of the laboratories: APC, AIM and IPGP. The project started in September 2012, and now are being officially in phase B. The IGOSat are being developed as a 3-unit cubesat, with scientific objectives are to measure the total electronic content (TEC) of the ionosphere by receiving GPS signals, and observing gamma ray in the inner Van Allen belt, or more specifically, the polar cups and the South Atlantic Anomaly. As come with other subsystems, the structure subsystem is now being designed in detail both internal and external layout of the satellite. The IGOSat’s structure is firstly based on the 3U cubesat skeleton from ISIS. However, due to the educating goal of the Project, the design of using the ISIS skeleton and building a new structure by students are implemented parallel. Although this new structure have some borrowed ideas from the ISIS’s one, this give students chances to learn and gain hand-on experiences by beginning from the first step of design. This thesis, which is re-organized from the technical report IGOSat Study of Mechanical Conception (STR-NT-02), is divided into four parts: Introduction: Introduces the scope of the thesis. I. Modeling IGOSat Shows the detail layouts of the IGOSat with different plans proposed. II. Mechanical testing proposals Proposes tests, which include definitions, scenario, and profiles to verify the IGOSat in launching conditions. III. Mechanical simulations Shows results of proposed tests simulation. IV. Conclusions Concludes the done works, unsolved problem and the success of the internship.

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I. MODELING IGOSAT In order to have a best design, the satellite has been being visualized in CATIA based on the model from the preliminary design concept.

1

1

1 12

2

12

2 3

2

4

4

4 5

5

5

6 3

3

7 8

9 10 3

7 6

9 10

7 6

3 3

3 11

8

Sketch-up model

11

3

3 13 9

11 CATIA model (Plan A)

CATIA model (Plan B)

Figure 1: IGOSAT CAD Models 1. Scintillator 2. Electrical Power Supply 3. Empty board 4. UHF/VHF antenna 5. Module AMSAT 6. Attitude determiner 7. On board computer 8. Sun sensor 9. GPS receiver 10. Actuator board 11. GPS antenna 12. EASIROC 13. Mass balancer

Basically, the satellite has two options while moving on orbit: The Z direction of the satellite points to Nadir, or coincident with the Astronomical Horizon. Consequently, it needs two plans to organize all subsystems and the adaptions of the structure. To have a minimum number of differences between these two plans, the modifications will be made on the first cube of the satellite (the cube contains the GPS payload). 2

1. Assembling IGOSAT With the benefit in model visualization that CATIA takes with, the satellite has been modeled impressively as a real model; and this CATIA model gives much information in order to detect contradictions between parts, mass distribution, etc. Subsequently, there appeared problems that need to be modified the Google Sketch-up model due to the new update of components and their more exacted shapes. At this time, there is lack of information to sketch up a CATIA model for the third cube of the satellite which contains the payload Scintillator and the Electrical Power Supply subsystem. Conversely, it seems to have enough information for drafting other cubes, including the UHF/VHF antenna. 1.1. UHF/VHF antenna To be consisted as much as possible with the old Google Sketch-up model from the Preliminary design, the antenna will be positioned in between the top and the middle cubes. This space, normally, it not used for placing big components such as antenna or module boards. Thus, the task is to find a solution in order to have an adaption. Fortunately, all ISIS products are compatible with each other, and it is much easier to use the ISIS deployable antenna, which has its mounting holes coincident with holes on the Aluminum structure ribs.

Figure2: Connection between antenna and structure

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1.2. Middle cube (second cube) The second cube provides space for the AMSAT communication board, the ISIS Onboard Computer, the ADCS board and an empty board. The empty board is used to assure the mass budget. In the Sketch-up model, the positons of these components in top-down direction are AMSAT board, ADCS board, Computer and the empty board is at the bottom. However, all of these three electrical modules: the AMSAT board, the ISIS computer and the ADCS board take extra-spaces of 4mm, 2.8mm and 6mm respectively, at the PCB’s bottom side. To prevent the intersections between the upper part of the ADCS board and the lower parts of the AMSAT board, between the air magnetorquer and the Computer, the Computer and the ADCS board will be switched their positons to each other; and the distances among all components in the cube are a little bit modified as shown in the figure 3.

Figure 3: Organization in the middle cube

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1.3. Bottom cube (first cube) 1.3.1. Plan A

The bottom cube will contain the OEM 615 GPS receiver, a Monitoring card and a Dual-frequency GPS antenna. In case of the ESP card and thermal sensors’ positions are well designed, there is no longer need for the Monitoring card. Then the first cube is a home for only the GPS payload.

Figure 4: The OEM 615 GPS board (right) and its adaption on PCB (left)

The mechanism to hold the GPS antenna is to pocket a hole in the center of one or two PCBs. It is reasonable due to the mechanical design of most antennas where fixations are near connectors. Moreover, all other empty PCBs will have one hole at the center to create connection space between the GPS antenna and the Receiver card. Each PCB is separated from other by 13mm spacers. The layout of the bottom cube is showed in the figure 5.

Figure 5: Bottom cube layout side view (left) and isometric view (right)

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1.3.2. Plan B

The plan B will be implemented in case the satellite points directly to the Nadir (along Z direction). This plan has the same organizations for the top and the middle cubes, where the directions of modules are not affected by the change in orbit of the satellite. Then first difference in comparing with plan A is the modification of the rib part (figure 7). The side frame-rib and rib-modules mounting holes are now on the same plane, not being perpendicular as in the plan A. It, therefore, leads to the decrease of PCB’s size to 85x85 mm; because the PCBs which satisfy the PC104 standard are now become too big. The cutting on the OEM 615 card’s PCB adaptor is made for the electrical connections of solar panels and the GPC payload.

Figure 6: The OEM 615 GPS board and its modified adaption on PCB

Figure 7: Modified rib

Figure 8 shows the layout of the bottom cube in the plan B, where the gravity center is located far from the geometrical center by 17mm, 0, 36 mm and 0 mm, along the longitudinal and transverse directions respectively. While the GPS antenna has weight of 185 g, the GPS receiver is only 25 g. to rebalance the cube, it is necessary to put an aluminum mass balancer. This balancer is, of course, has the 85x85 mm size, 10 mm of thickness, and has a hole of 30 mm diameter at its center. The 21 mm distance between the balancer and the OEM 615’s PCB adaptor also gives enough space for the card Receiver to connect to other modules by itself via cables.

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Figure 8: Bottom cube plan B’s layout: side view (left) and isometric view (right)

2. Assembly procedure Due to the structure is designed base on the ISIS structure, it is better to implement the same assembly procedure with one form ISIS. Each cube will be assembled with its components and modules first. The top and the middle cubes will be rigid together by connecting them with the ISIS UHF/VHF antenna; and then after, all three cubes will be mated with side frames where kill switches have already fixed on. Further detail on the procedure can be found in the reference [R03], and the first IGOSat STM Integration Manual [A06].

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3. IGOSat skeleton The simplified skeleton is created based on these following criteria: -

-

-

Similarity to the ISIS structure: the new skeleton is designed based on the inspiration from the ISIS’s product. It was created as simplest as possible, such as all fillets and chamfers have radius below 2mm are removed or replaced; increase thickness despite of increasing mass and volume. However, the most essential criterion is to guarantee the adaption of the new skeleton (simulation proved). Homemade ability: The more the skeleton is made at the APC, the easier situation we have. Not only the price is advantageous in this case, but we can also learn, and experience to produce our own structure. The side frame (as called in ISIS structure) is divided into two smaller parts after simplifying. Then the rib parts and one part of side frame (called side-frame-B) can be produced at the APC. Price: Aluminum 7075 or 6061 are high quality and expensive. It will take time to order some of them. Then the Aluminum 2017, which has nearly the same mechanical properties with the required ones, was used instead, and was adequate for the STM tests.

Figure 9: IGOSat Skeleton

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II. MECHANICAL & ENVIRONMENTAL TESTING PROPOSAL 1. Purposes In every production, tests are intended to validate and demonstrate hardware and software design concepts that can satisfy requirements, and to assist in the evolution of designs from beginning phase to the operational phase. The tests developed in this project will be used to identify problems early in the design evolution, and then perform as criteria for the correction. Basically, the satellite must pass through three phases of testing, includes Qualification tests, Acceptance tests, and Pre-launch tests, before operate its missions in space. Almost space vehicles need to be validated at every level, from single units to sub-system and then to the whole satellite, by the following tests: Mechanical tests (Acceleration (Quasi-Static), Resonance survey, Sinusoidal vibration, Random vibration, Shock loads), Environmental tests (Thermal vacuum, and Thermal cycling), and Functional tests such as Electromagnetic compatibility. Because of the thrust and roll motion during launching, the satellite will be accelerated and suffered the force made due to the small values of excitation frequencies compared to the fundamental frequencies of the launching vehicle. Acceleration test or Quasi-static loads test is made for the purpose of examine the stiffness, the strength adequacy of the primary structure, the critical structural interfaces between cubesat and launcher, cubesat and its modules. Resonance survey or Natural frequencies test is a pre-requested criterion for other vibration tests. This inspection is essential to find which frequencies can lead to the resonance of the system, and from its results, designers can take modifications to models meet requirements. In addition to the Quasi-static Loads test, Sinusoidal vibration test at lowfrequencies (less than 50 Hz) demonstrates the ability of the satellite to survive or to operate in the sinusoidal transient, in sinusoidal environment, or in order to detect material, connections or parts that might not adequate to the real environment. Not only act as a support test, Random vibration test is the most essential in the qualification process. It assesses units subjected to the random dynamic environment, which might be experienced during flight. Shock test demonstrates the capability of the satellite structure and instruments to meet the requirements of strength, stiffness or both, under the loads conducted when launching, or separating from rocket, at the burning up stages of the rocket or deployment stage of the satellite. Thermal vacuum test will show how good the system faces under vacuum conditions and temperature extremes. Thermal cycling test is to prove the ability of the system to meet the deign requirement in thermal stressing environment. 9

Basically, two levels of these tests are specified include: the qualification test and the acceptance test. In comparing with the qualification test, the levels of the acceptance tests are 3dB lower for vibration tests and are 80% for the others. Orderly, tests will be conducted from Structure and Thermal Models (STMs), Engineering Models (EMs) to Qualification Models (QMs) and Flight Models (FMs). The STMs is responsible for mechanical and thermal behavior, while the EMs undergo electromagnetic compatibility and other functionality tests. The QMs and FMs can be understood as combinations of STM and EM tests at level of qualification and acceptance. 2. Scenario As useful and essential as the NASA GEVS GSFC 7000 Standard, the Military 1540B Standard from US Defense and the VEGA Launch Vehicle Program General Specification are the referred document to describe the ways how to implement the tests. 2.1. Sinusoidal vibration test The satellite will be attached to a vibration fixture by a flight adapter which is the P-POD or ISIS-POD in this case. The force of vibration will be applied on each of three axes while all deployable components such as antenna, solar panels must be packed as in the launching conditions. During the testing time, instruments for detecting the performance of the satellite can be installed to measure the vibrating responses on the three axes. 2.2. Random vibration test This test will be implemented as same as the sinusoidal test. However, instead of using predictable frequencies and amplitudes, the profile for this test includes random values. 2.3. Shock test For this test, the satellite will be place inside the orbital deployer, or otherwise supported to prevent re-contact between separated parts thereof. In cases of significant shocks from subsystems not on board the satellite under test, such as launch vehicle separation, or burning up stages, the simulation adapter will simulate these imposed forces.

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2.4. Thermal vacuum test The satellite with all of its modules will be placed in a thermal vacuum chamber and the functional test will perform to assure the readiness of the system. During the period of reducing pressure to the specified level, all equipment and modules will not have electrical power supply, as same as in real launching, and to satisfy the specification requirement for cubesats. The temperature cycle will begin at ambient temperature, then be reduced to the specified low level and stabilized. From the lowest level, temperature will be increased to the highest specified level, stabilized, and returned to the ambient level in order to complete one cycle. The levels of tested temperature shall be at least +/-10 Celsius degree above/below the respective maximum/minimum predicted flight. During each cycle, temperature stabilized components shall be active, and measurement instruments can be installed to calculate result as expected to the real conditions. *: Ambient environment or room condition is defined as with temperature 23 (+/-10) Celsius degree, atmospheric pressure 101 (+2/-23) kilopascals and relative humidity of 50 (+/-30) percent. 2.5. Thermal cycling test The basic difference between the thermal vacuum test and the thermal cycling test is the pressure level. During the thermal cycling test, the cubesat will be placed in a thermal chamber at ambient pressure, and unfavorable combinations of temperature and humidity shall be prevented to assure there is no moisture deposition. The thermal cycles will only begin when the relative humidity of the inside spaces of the satellite is below at which the coldest temperature can cause condensation. 3. Test profile Due to the limit of referenced documents and time, at first steps, the profile for each test will be based on the Specifications Documents of the VEGA Rocket from European Space Agency (ESA), which is assumed as the Launch vehicle in this Project, the System Requirements Document for QB50 projects, and referred from the Mission Design for the OUFTI-1 Cubesat. However, the final profiles must be constructed and improved base on the specific mission requirements of the Project and the exact Launch Vehicle service. 3.1. Acceleration (quasi-static) test The Cubesat shall be tested in all three directions X, Y and Z in 1 minute with the amplitudes 13g and 10.8g respectively for Qualification test and Protoflight test.

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3.2. Resonance survey The system and all of its subsystems shall pass the test with the lowest natural frequencies higher than 90Hz. 3.3. Sinusoidal vibration test The maximum predicted sinusoidal vibration is over a frequency range of 20 to 2000 Hz for flight excitation, and 0.3 to 300 Hz for ground transportation excitation. The test duration on each axis shall be three times the expected flight duration and the detail is in table 1. 3.4. Random vibration test The random vibration environment imposed on the satellite during the lift-off by aerodynamic excitations and transmitted structure-borne vibration. The maximum predicted values are defined as a power spectral density in range of 20 to 2000 Hz, with a frequency resolution of 1/3 octave. Test profile is in table 2. Table 1: QB50 Sinusoidal vibration test characteristics

Method Sine vibration test Reference frame Direction Sweep rate Profile

Method Random vibration test Reference frame Direction RMS acceleration Duration Profile

Qualification

Acceptance FEM simulation + Test

Protoflight

Required

Required

Required

BRF

BRF

BRF

X, Y, Z 2 oct/min Frequency Amplitude (Hz) (g) 5-100 2.5 100-125 1.25

X, Y, Z 4 oct/min Frequency Amplitude (Hz) (g) 5-100 2 100-125 1

X,Y,Z 4 oct/min Frequency Amplitude (Hz) (g) 5-100 2.5 100-125 1.25

Table 2: QB50 Random vibration test characteristics Qualification Acceptance FEM simulation + Test

Protoflight

Required

Required

Required

BRF

BRF

BRF

X, Y, Z

X, Y, Z

X, Y, Z

8.03 g

6.5 g

8.03 g

120 s Frequency Amplitude (Hz) (g2/Hz) 20 0.01125

Frequency (Hz) 20

60 s Amplitude (g2/Hz) 0.007

120 s Frequency Amplitude (Hz) (g2/Hz) 20 0.01125

12

130 800 2000

0.05625 0.05625 0.015

50 200 640 2000

0.007 0.035 0.035 0.010

130 800 2000

0.05625 0.05625 0.015

3.5. Shock test The maximum predicted shock environment is specified as a maximum absolute shock response spectrum determined by the response of a number of single degree of freedom systems using Q = 10. The Q is the acceleration amplification factor at the resonant frequency for lightly damped system. The frequency range is from 10 to 10000 Hz, with intervals of 1/3 octave. Detail profile is presented in table 3. Table 3: QB50 Shock test characteristics Qualification Method Shock test Reference frame Direction Q factor Number of shocks Profile

Acceptance Test

Qualification

Required

Required

BRF

BRF

X, Y, Z 10

X, Y, Z 10

2

2

Frequency Amplitude (Hz) (g) 30 5 100 100 700 1500 1000 2400 1500 4000 5000 4000 10000 2000

Not required

Frequency (Hz) 100 1000 2000 5000

Amplitude (g) 30 700 700 400

3.6. Thermal tests These tests at first steps are based on specifications of the IGOSAT’s components, referred from the VEGA Requirement Specifications and the OUFTI-1 Satellite (table 4 and table 5). Table 4: Thermal vacuum test characteristics Qualification Method Atmospheric pressure Relative humidity Temperature range Change rate (heating) Change rate (cooling)

Acceptance TBD

1 mPa TBD [-81,218] < 20 oC/min (internal), > 20 oC/min (external) 2 ÷ 3 oC/min

[-65,175] < 20 oC/min (internal), > 20 oC/min (external) 2 ÷ 3 oC/min

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Number of cycle Duration at boundaries

40 (for electronic equip) 10 (for non-electronic equip) 2h

4 1.6h

Table 5: Thermal cycling test characteristics Qualification Method Atmospheric pressure Relative humidity Temperature range Change rate (heating) Change rate (cooling) Number of cycle Duration at boundaries Stabilization criterion

Acceptance TBD

TBD TBD [-81,218] < 20 oC/min (internal), > 20 oC/min (external) 2 ÷ 3 oC/min 10 2h 1 oC/h

Covered by the Thermal vacuum test

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III. SIMULATIONS RESULTS & DICUSSIONS In this project, there are components including the ADCS board, the Power Supply Subsystem, Scintillator Payload and the Adapter for GPS antenna, which is developed by students and interns. Therefore, not only the whole satellite, but these subsystems must be also verified by all tests before performing missions. In the mechanical point of view, these components will undergo environmental tests specified in requirements for every cubesats. Modeling and Meshing Due to the complexity of the system, there will be a huge number of calculations when simulating. Therefore, it is better to analyze a new simpler model derived the detail model. In the new model, all electrical devices include resistors, capacitors, micro-controllers, connectors, etc., some not essential holes, surfaces, screws, nuts will be removed. The goal is to simulate the satellite in basic shapes model which is able to mesh quickly and run analyses faster but small deviated results in comparing with the real complex model. The meshing element sizes for all models using in simulation are tetrahedral automatically for parts which are not to have high precision calculations. At parts which need to be analyzed in detail, the meshing type is hex dominant meshing. To be noted that, all tables in the Appendix D summarize only the maximum results of each simulation, in each direction. Photos are taken from some of these results do not represent for the final conclusions. There will be some illogical results because areas affected most in different directions are different.

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Boundary conditions Commonly, each subsystem is represented by a PCB which satisfies the PC104 standard. At the moment of implement first simulations, it is simpler to fix all PCBs at their mounting holes (inner faces) in simulator for vibrations tests and shock test. These fixations will be used at all three impact directions (X, Y and Z directions). Test profiles are showed in part II. For quasi-static test, the fixations are depended on the directions of applied forces. Thus, in each direction, the outermost face of the PCB will be the fixation for the test. However, in Z direction, due to the PCB’s thickness is only 1.6 mm, the fixed supports will be at inner faces of 4 mounting holes. The force applied is calculated from the mass of each components multiplied with 13 times of acceleration (F = m*13*G = 127.53*m). In order to simulate the satellite inside the POD, it is necessary to simulate the fixations where the IGOSat contact with the POD. Basically, the top, bottom and two lateral faces of the structure’s rails will contact with the rails of the POD. It leads to two ways to set up the boundary conditions: fixed 8 lateral faces of the four rails; or fixed four bottom faces and let four top faces have ability to displace in Z direction. The second way was chosen due to these following reasons: The cubesat will be pushed out by POD’s deploying spring; there are some tiny gap between rails of the cubesat and rails of the POD due to some acceptable tolerances; it is easier to see how all lateral faces of the cubesat react in simulation without any fixation. The fixations set on each single cube are at their four mounting holes with the structure. 1. Simulating GPS adapter Obviously, the way how the adapter deforms, or breaks is not much different between versions of one and two layers of PCB. However, the levels of effect on these versions are convincible. The maximum deformation at quasi-static simulation decreases 6.5 times from 0.654 mm to 0.1 mm, at 1 layer PCB version and 2 layers PCB version respectively. Moreover, the natural frequencies at this level of simulation are still acceptable for both versions, but it is considerably different at assembly level (see Simulation bottom cube).

Figure 10: Acceleration deformation GPS adapter.

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Table 6: Simulating GPS adapter of 1 PCB summary

X axis

Y axis

Z axis

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Acceleration test

Random vibration test

Sinusoidal vibration test

0.15

0.104

0.020

11.67

66.22

1.56

0.15

0.0654

0.020

11.67

37.15

0.87

0.654

0.654

0.021

19.14

24.02

0.96

Photo

2. Simulating Skeleton Comparing results from two cases simulating the skeleton (table D-4 and D-5), it can be figured out that the model with carbon fiber plates has smaller deformations and is affected under lower stresses although, in quasi-static simulations, the force’s magnitude depends on the mass of components. The forces applied on models with carbon plates (0.523 kg) and without these plates (0.218 kg) are 66.7 N and 27.8 N respectively, but both deformations are tiny when scaling with the structure’s dimensions. The stresses in random vibration simulations have considerably differences in between models with and without carbon plates. While they increased 2 times stress on model with carbon plates in Z direction, these other directions had inverse increments (7.4 and 4.74 times in X and Y directions). Nevertheless, all stresses calculated are in elastic deformation range for Aluminum 7075-T6 and Carbon fiber materials (Appendix C).

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Figure 11: Deformation (maximum) result of acceleration simulation of skeleton

Figure 12: Deformation (maximum) result of acceleration simulation of skeleton with carbon fiber plates

Table 7: Simulating GPS adapter of 2 PCBs summary

X axis

Y axis

Z axis

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Acceleration test

Random vibration test

Sinusoidal vibration test

0.0149

0.070

0.020

12.19

115.77

1.28

0.0149

0.077

0.020

12.19

64.5

0.72

0.100

0.370

0.038

8.74

32.15

1.59

Photo

18

Table 8: Simulating skeleton without carbon plates summary

X axis

Y axis

Z axis

Acceleration test

Random vibration test

Sinusoidal vibration test

0.0042

0.180

2.05E-5

2.11

65.71

0.0045

0.0016

0.12

2E-5

0.79

35

0.003

0.002

1.4E-4

2E-5

0.99

0.054

5.64E-7

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Photo

3. Simulating Bottom cube The model using one PCB to hole the GPS antenna has been found to have one natural frequency below 90 Hz (79.908 Hz). Then the design has been modified to use two PCBs to hold the antenna which has the minimum simulated resonant frequency is 203.8 Hz.

19

4. Simulating Top cube These simulations were done by simplifying each group of following components into one singular part: The Scintillators and SiPM, the four Lithium batteries, the four battery brackets. The simplified Scintillator is assumed to be the full 13x13x16.35 mm cube of LaBr3, the more massive material. Table 9: Simulating skeleton with carbon plates summary

X axis

Y axis

Z axis

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Acceleration test

Random vibration test

Sinusoidal vibration test

0.0024

0.017

1.6E-4

1.16

8.83

0.12

0.002

0.0176

2E-5

1

7.38

0.15

0.0014

2E-4

2E-5

1.6

0.11

1.78E-6

Photo

20

Figure 13: Deformation (maximum) result of acceleration simulation of bottom cube

Figure 14: Deformation (maximum) result of random vibration simulation of bottom cube

Table 10: Simulating Bottom cube summary

X axis

Y axis

Z axis

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Acceleration test

Random vibration test

Sinusoidal vibration test

0.044

0.032

2E-5

58.19

4.26

1E-4

0.017

0.0087

1.98E-5

30.82

0.41

1E-5

0.082

0.434

3.25E-5

20.60

95.49

0.003

Photo

21

Figure 15: Deformation (maximum) acceleration simulation of top cube

result

of

Figure 16: Deformation (maximum) result of random vibration simulation of top cube

Table 11: Simulating Top cube summary

X axis

Y axis

Z axis

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Maximum VonMises stress (MPa)

Acceleration test

Random vibration test

Sinusoidal vibration test

0.955

0.181

2.37E-5

28.39

26.36

5.5E-4

0.0642

0.243

2.40E-5

20.78

42.76

6.9E-4

0.165

0.976

7.416E-5

32.01

223.08

0.006

Photo

One remarkable result is the 0.95 mm deformation of the adapter board of the Scintillator. Although the Von-mises stress is still inside acceptable range, this result 22

could be worse when the test is applied at whole satellite level. On the other hand, only fix the Scintillator at its bottom cannot guarantee the strong of the mechanism while inside components of the Scintillator are still not simulated. It, therefore, is proposed to have another adapter on the top of the payload, to form a “Sandwich” mechanism. 5. Simulation the whole satellite Unfortunately, due to the limit of the computer’s memory, it is apparently impossible to launch the simulation for the whole model of the IGOSat. It, therefore, was simplified by replacing all subsystems by greatly simple shapes: rectangular, cylinder, etc. Nevertheless, on the other hand, the new simplified model is impressively as same as the real STM model. Then it is good to compare the simulation results with the real test results. It is easy to figure out the weakness area on the IGOSat, from the deformation animations, is the Scintillator payload. (Summarized at table D-8). The adaption for the GPS antenna needs to be considered carefully also. On the other hand, although the weight of this payload was increased from 38gr to 76 gr, the deformation and stress appeared is still decreased. Then the ability of the “Sandwich” mechanism is confirmed.

Figure 17: Acceleration simulation on Z axis

The random vibration is always the most damage cause during mechanical tests or launching period. Figure 19 and 20 show the response of the IGOSat on three axes X, Y, Z in Random vibration simulation. The responses on all three axes vibrate significantly, or energy intense, from 700Hz to 1100Hz although the magnitude on each axis is different. There are also some smaller peaks around 400-500Hz appear on Z and Y axes. This phenomenon consists with the range of natural frequencies, which is showed in figure 18 below.

23

1400 Deformations

Deformations (mm)

1200

Average deformation

1000 800 600 435,35772

400 200 0

Frequencies (Hz)

100

1000

Figure 18: IGOSat simulated natural frequencies (first 50 modes).

A (red): Test profile B (green): PSD on Z axis C (blue): PSD on X axis D (purple): PSD on Y axis

Figure 19: Random vibration's Power spectrum densities on three axes

24

Figure 20: Random vibration's Power spectrum density on Z axis

Table 12: Simulating simplified IGOSat

Maximum Deformation (mm) Maximum VonMises stress (MPa) Maximum Deformation (mm) Y axis Maximum VonMises stress (MPa) Maximum Deformation (mm) Z axis Maximum VonMises stress (MPa) X axis

Acceleration test

Random vibration test

Sinusoidal vibration test

0.0127

0.159

0.019

12.029

118.07

14.534

0.0254

0.55

0.020

13.352

61.58

14.870

0.121

0.768

0.045

16.027

141.33

3.327

Photo

25

IV. CONCLUSIONS AND PERSPECTIVES At the end of this internship, several aspects of the mechanical conception of the IGOSat were swept through the collaboration among the satellite’s subsystems, and advices from board of supervision. In the first part, all general requirements and specific constrains of the IGOSat were summarized. It helps to highlight the strictly criteria, creation needs, and to contribute the roadmap for the design process. From this, the 3D visualization can be implemented. Continuing the previous work of Pedro Lopes (2015), the visualization of the IGOSat’s layouts was finished by using CATIA, and basic verified by using ANSYS. Particularly, the structure subsystem, which is inspired from ISIS products, has been customized to be simpler and able to be produced by the IGOSat team. The next step has been done is to study mechanical qualification process. As usual, before operating in real life, a product must pass through all testing requirements. For the IGOSat, although the space environment is critical, period during launching cause most mechanical damage. Then, the proposal for the mechanical test was made by applying popular standards: NASA GSFC 7000, QB50 Recommendations, and the assumption of using VEGA launching service. The first IGOSat’s STM was also built up, and was verified by vibration tests. Apparently, there are some problems remained and unsolved, but it gives chances to gain hand-on experiences for qualification process, and ideas for future corrections and improvements.

26

REFERENCE [R01] P. Lopes, “Preliminary Design and Thermal Study of IGOSat Project”, Master thesis, Université Paris Diderot, August 2014. [R02] A. Denis, C. Asma, C. Bernal, R. Chaudery, Z. de Groot, J. Guo, D. Kataria, D. Masutti, R. Reinhard, M. Richard, T. Scholz, G. Shirville, F. Singarayar, B. Taylor, P. Testani, J. Thoemel, and W. Weggelaar “QB 50, System Requirements and Recommendations”, Issued 7, 13 Feb 2015. [R03] “Assembly Manual ISIS 2-Unit Cubesat Structure”, 20 Dec 2012, available on http://www.isispace.nl/cms/ [R04] US Defense, Military Standard, “Test Requirements for Launch Upper-stage, and Space Vehicles”, 10 Oct 1982. [R05] S. Galli, “Mission Design for the Cubesat OUFTI-1”, Master thesis, University of Liège, 2008. [R06] J. Thiéry, “VEGA User’s Manual”, Issued 04, April 2014, available on http://www.arianespace.com/index/index.asp [R07] S. L. Grelle, “Analysis system”, 2015

27

Appendix A:

STRUCTURAL REQUIREMENTS

Fundamental requirements The external dimensions of the 3U IGOSat are of 10 ±0.01 x 10 ±0.01 x 34.05 ±0.01 cm. The space inside each cube has the dimensions of 9.6 ±0.01 x 9.6 ±0.01 x 7.5 FR 2: ±0.01 cm. The total mass of the cubesat is expected as 3.75 kg, include the margin of FR 3: 30% (S. L. GRELLE, 2015) FR 1:

FR 4:

The assembly’s direction is on the Z direction and centered at the origin of the coordinate system.

FR 5:

Components cannot exceed 9 mm outside the 10 cm cube by using ISIPOD , but 6.5 mm outside is expected.

FR 6:

The contact rails between deployer and the satellite must have a minimum width of 8.5 mm (CDS 3.2.5).

FR 7:

Aluminum 7075, 6061, 5005, and/or 5052 will be used for both the main CubeSat structure and the rails (CDS 3.2.15).

The CubeSat rails and standoff, which contact the P-POD rails and adjacent FR 8: CubeSat standoffs, shall be hard anodized aluminum to prevent any cold welding within the PPOD (CDS 3.2.16) FR 9:

The center of gravity must be within a sphere of 20 mm from the satellite’s geometrical center, but can be extended to 70mm (CDS 3.2.14)

Payloads requirements  Scintillator: Dimensions: TBD Mass: 0.08 kg (+ 20%) Requirements of spacing and shielding [R01]: - Need as few obstructions as possible - Operational temperature range in between 0 and 40 Celsius degree - Particles from lateral and bottom sides must be avoided  GPS receiver: Dimensions: - GPS Receiver: 7.11*4.57*1.18 cm - GPS Antenna: 6.65(dia.) * 4.9 (h.) cm Mass: - GPS Receiver: 24 g - GPS Antenna: 185 g Requirements of spacing and shielding [R01]: A-1

-

The antenna must be placed at the back of the satellite in order to point to the GPS satellite.

Deployment requirements UHF/VHF Antenna: The antenna will be positioned between the top and the middle cubes, and therefore, an adaption of mechanical support is needed. Kill switches: The two kill switches will be put at the bottom cube, belonged to the side frames (one switch for each frame). The preliminary design can be as same as the design in ISIS structure. Adaptation for small modules The components including OEM 615 GPS Receiver card, Dual frequencies GPS antenna, AMSAT Communication board, will be held by PCBs which has size afforded to the cubesats dimensions. They will connect with other components through adaptions made on PCBs. Due to unstandardized dimensions; the positions for these components need to be chosen by criteria of saving space and number of modifications on other cubesat-size parts. Sun sensors will be place in space between the middle and the third cube in order to have sensing holes face out to the Sun. Thus, they need adaptions also; and the idea is create a 4-walls band as same as the ADACS Payload Walls from Cubesatkit.

ADACS payloads wall (Source: Cubsatkit.com)

Mass budget The total mass of the satellite must not exceed 4 kg, including their margins. The organization inside the cubesat must assure the distribution of mass, in order to have the distance between the center of gravity and the geometrical center less than or equal to 2 cm.

A-2

Summary table Table A-1: Mass budget summary of sub-systems No.

Sub-system

Component Name

Dimension (cm)

Quantity

Mass (g)

Mass (plus 30% margin)

IGOSAT

TBD

01

TBD

(g) TBD

Novatel OEM615, Antenna TW 3802

7.11*4.57*1.18, 4.9*6.65

01 01

24 185

31.2 240.5

Onboard Computer

ISIS Onboard Computer

9.6*9.0*1.24

01

95

123.5

4

ADCS

IGOSAT ADCS board Sun sensor [xxx] Magnetometer [xxx]

9.589*9.017*2 3.3*1.1*0.6 TBD

01 05 01

193 5 TBD

250.9 6.5 TBD

5

Communication

ISIS Antenna, AMSAT-F TT&C Card

9.8*9.8*0.7 9.589*9.017*2.7

01 01

94 194

122.2 252.2

6

Electrical Power Supply

IGOSAT ESP board [xxx] Solar panel

9.8*8.26*6.0 Various

01 13

TBD 65

TBD 84.5

7

Mechanical Structure

ISIS Structure (include screws, fasteners) Carbon fiber plates Supports

10*10*34.05 Various Various

01 5 N/A

226 302 N/A

293.8 392.6 N/A

>2183

>2838

1

Scintillator

2

GPS

3

Total mass: Further detail is presented on Appendix B.

A-3

Appendix B:

Item name Heater

Skeleton Side frame Rib 800 Rib 737 top Rib 737 bottom Screw M2.5x6 Fastener Carbon fiber covering 1U top plate 3U plate for rib 3U plate for side frame Screw M2.5x8 Cube 1 Stack Spacer Washer M3x0.5 Bus spacer M3x12 Hex Nut M3x12.4 Screw M3x8 Threaded rod Item name Hex Nut M3x8

List of structure components

Item code

Quant ity

General dimensions (cm)

Mass (g)

Material

Compare with ISIS structure

STM 1

Not defined

STM 2 STM 2.1

10*10*34.05

STM 2.2 STM 2.3 STM 2.4 STM 2.5 STM 2.6 STM 3 STM 3.1 STM 3.2 STM 3.3 STM 3.4

226

2

34.05*10*15.2

69,43

6

8.68*1.17*1.09

6,76

3

8.28*1.17*1.09

6,67

3

8.28*1.17*1.09

6,67

24

0.25*0.6

0,27

24

0.3*0.15

N/A

Aluminum 7075 Aluminum 7075 Aluminum 7075 Aluminum 7075 Stainless steel Stainless steel

Same Modified Modified Modified Same Same

Spring

302

STM 4 STM 4.1 STM 4.2 STM 4.3 STM 4.4 STM 4.5 Item code

Note

1

9.8*9.8*0.2

23

2

32.5*8.26*0.2

67

2

32.5*8.26*0.2

67

52

0.25*0.6

0,22

Carbon fiber Carbon fiber Carbon fiber Stainless steel

N/A N/A N/A

Mass density 1800 kg/m3

N/A

>40 40

0,08

4

0.3*0.05 & 0.5*0.05 0.3*1.2 & 0.5*1.2 0.3*1.0

4

0.3*0.8

0,44

4

0.3*7.5

3

16

Quant ity

STM 4

4

0,84 0,72

General dimensions (cm)

Mass (g)

0.3*0.8

0,52

Stainless steel Stainless steel Stainless steel Stainless steel Stainless steel Material Stainless steel

Insidehole & Outside cover

Compare with ISIS structure

Note For OEMB-1

. 6 Screw M3x4 Empty board

STM 4.7 STM 4.8

Cube 2 Stack Spacer

STM 5

Washer M3x0.5

STM 5.1

Bus spacer M3x12 Bus spacer M3x18.4 Hex Nut M3x5.2

STM 5.2 STM 5.3 STM 5.4 STM 5.5 STM 5.6 STM 5.7

Screw M3x8 Threaded rod Empty board

Item name Cube 3 Stack Spacer Washer M3x0.5 Bus spacer M3x? Bus spacer M3x? Hex Nut M3x? Screw M3x8 Threaded rod

Item code STM 5 STM 6.1 STM 6.3 STM 6.4 STM 6.5 STM 6.6 STM 6.7

Item name

Item code

Antenna adapter

STM 7

615 adaption 8

0.3*0.4

0,27

Stainless steel

5

9.017*9.017*0.1 6

TBD

TBD

>33

72

12 2

0.3*0.05 & 0.5*0.05 0.3*1.2 & 0.5*1.2 0.3*1.84 & 0.5*1.84

0,08

0,84 1

Stainless steel Stainless steel Stainless steel Stainless steel Stainless steel

4

0.3*0.52

0,325

4

0.3*0.8

0,44

4

0.3*7.5

3

1

9.017*9.017*0.1 6

TBD

TBD

Quant ity

General dimensions (cm)

Mass (g)

Material

0.3*0.05 & 0.5*0.05 0.3*4.0 & 0.5*4.0 0.3*0.3 & 0.5*0.3

TBD

4

0.3*1.8

TBD

4

0.3*0.8 & 0.5*0.8

0,44

4

0.3*7.5

36 4 4

Quant ity

General dimensions (cm)

0,08

TBD

3

Mass (g)

Insidehole & Outside cover

Stainless steel

Depends on AMSAT module

Compare with ISIS structure

Stainless steel Stainless steel Stainless steel Stainless steel Stainless steel Stainless steel

Material

Note

Detail length depends on Scintillat or

Compare with ISIS structure

Note

10,72

B-2

Hex Nut M2.5 Washer 2.5 Screw M2.5x8 Threaded rod Kill switch adapter Bracket A Bracket B Screw M1.6x4 Hex nut M3 Pin String Foot

STM 7.1 STM 7.2 STM 7.3 STM 7.4 STM 8 STM 8.1 STM 8.2 STM 8.3 STM 8.4 STM 8.5 STM 8.6 STM 8.7

8

0.25*0.42

0,16

8

0.3*0.05

0,08

8

0.25*0.8

0,3

4

0.3*7.5

3

Stainless steel Stainless steel Stainless steel Stainless steel

4,2 2

2.1*0.75*0.37

0,225

Aluminum

Same

2

2.1*0.75*0.06

0,226

Aluminum

Same

4

0.16*0.4

0,07

4

0.16*0.13

0,06

2

0.34*0.34*2.6

0,7

2

4.5*1.75

0,2

2

0.55*1.15

0,5

Stainless steel Stainless steel Stainless steel Stainless steel Stainless steel

Same Same Same Same Same

B-3

Appendix C:

Simulation Materials Properties

Density (kg/m3)

Modulus of Elasticity (GPa)

Poisson’s ratio

Aluminum 7075-T6

2810

72

0.33

Carbon fiber reinforced polymer

1800

181

0.1

Copper

8900

110

0.37

Epoxy FR4

1850

3

0.35

Ferrite

5000

180

0.28

Stainless steel

7800

200

0.28

Material

(Source: Internet)

C-1