Structural Responses of a Conceptual Microsatellite Structure Incorporating Perforation Patterns to Dynamic Launch Loads
Abstract
:1. Introduction
- Pre-launch phase: This phase starts at the satellite’s assembly facility, continues at the test facilities that clear it for integration onto the launcher and to be transported to the launch facility (by land, sea, or air transport), and ends when the satellite is mated to the launch vehicle in preparation for launch.
- Launch phase: This phase typically lasts for approximately eight minutes, which is the time the launcher needs to propel itself and its satellite payloads from the launch site on the earth’s surface up to orbital altitudes. However, during these minutes, the satellite will experience the most intense dynamic loading it will ever experience during its full operational life, since during this phase it will be mechanically mated to the launcher. The launcher will convey to the satellite, through the launcher–satellite interface, the loading effects of multiple sources of dynamic excitation that all act on the satellite at a number of stages of the launch process.
- In-orbit phase: After the end of the launch phase at which the launcher will place the satellite into its orbit, the satellite will be floating in space, and hence it becomes a free body, from a mechanical point of view, without any points of support. This means that any outside forces acting on the satellite will lead to it moving as a full body, with no relative motion between its parts. Hence, no significant stresses will develop from this motion, and this phase is the least load-intensive part of the satellite’s operational life. However, some small dynamic loading will develop, due to moving parts such as momentum wheels, but their loads will be small compared to those resulting from launch or prelaunch load sources. In addition, some small deflections and resultant thermal stresses will arise due to thermal gradients between the sunlit and non-sunlit parts of the body, but their magnitudes are also small compared to stresses resulting from launch and prelaunch load sources.
- Quasi-static loads: Develop as a result of the response of the inertia of subsystem components and structural components to steady acceleration or slowly varying forces. These loads impose frequencies that are far from the natural frequencies of the components or the system; hence, this type of load will not induce any significant dynamic response. The loads are given in units of g (multiples of the earth’s acceleration constant g, 9.806 m/s2).
- Random loads: Develop as a result of non-deterministic loading sources, including the mechanical effects of acoustic loads, boundary-layer turbulence mechanical effects, high-frequency engine thrust oscillations, aerodynamic buffeting effects on the launcher fairing, sound pressure effects imposed in the satellite from launcher sources, and others. Random loads are characterized using acceleration power spectral densities (PSD), measured in units of (g2/Hz).
- Shock loads: Develop as a result of very short duration, high-intensity loadings, such as stage separations in multi-stage launchers, and especially the process of separation of the satellite from the launcher at the point of orbital injection. Random loads are characterized by a shock response spectrum (SRS).
2. Materials and Methods
2.1. Description of the Baseline Structure’s Component Geometry
- A Lightband Mark II separation ring [48], the interface between the satellite and the launch vehicle.
- A set of structural components shown, and listed, in Figure 2. The material for all these components was aluminum 6061 alloy, making this material dominant in terms of the design.
- The structural components listed above were interconnected through hex cap head screws of various lengths. The material selected for all fasteners was titanium 6Al4V. This material was selected for two reasons. The first was that its shear modulus is higher than that for aluminum 6061 alloy, hence leading to a higher resistance to shear. The second was its lower coefficient of thermal expansions relative to that for aluminum 6061, leading to superior thermal creep resistance. Knowing that thermal creep results from the thermal cycling that occurs in orbit, due to the entry and exit of the satellite into and from the shadow of the earth, thermal creep in the fasteners would have resulted in a loosening of the fasteners, which would have led to a reduction of the overall stiffness of the structure.
- To further ensure that the fasteners would not become loose due to vibrations and thermal creep, during both the launch and in-orbit operational phases of the lifetime of the satellite, they were locked into place using Heli-Coil inserts, as described in works by LaRocca et al. [49] and Rainville et al. [50], and others, which were selected based on the sizes defined for each set of fasteners.
2.2. Perforation Pattern Implementation
2.3. Descriptions of the Dynamic Launch Loads
2.3.1. Quasi-Static Loading Analyses
- The design factor is the product of the multiplication of two factors:
- The axial and lateral design limit loads, including the calculated design factor (DF), were:ADLL = 98.1 × 1.44 = 141.264 m/s2LDLL = 98.1 × 1.44 = 141.264 m/s2
- The design yield load was computed through multiplying the design limit loads by the yield factor of safety (FOSY). These loads were applied to the model, and the resulting stresses are known as the design yield stress. The value of FOSY was taken as 1.25 from the ECSS standard above, to afford a higher confidence in the capability of the design to withstand all the loads that it will undergo.Axial Yield Design Load = ADLL × FOSY = 141.264 × 1.25 = 176.58 m/s2Lateral Yield Design Load = LDLL × FOSY = 141.264 × 1.25 = 176.58 m/s2
2.3.2. Random Launch Loads’ Analysis
2.3.3. Shock Launch Loads’ Analysis
2.4. Methodology for Dynamic Loads’ Computational Analysis
- Solid models of the structural components, previously built using 3D modeling software, were imported into a new ANSYS Workbench project, within the Geometry ANSYS system.
- The ANSYS Geometry system data were then shared to two ANSYS modal analysis systems. The first analysis system was set to compute modal characteristics up to a frequency of 2000 Hz. The modal results from this analysis were fed into the subsequent random analysis stages, and the second analysis system was set to compute modal characteristics up to a frequency of 10,000 Hz, to feed its results to the subsequent response spectrum stages. The ANSYS modal analysis systems were required because both the random and shock analyses depend upon precomputed modal results, as part of the modal superposition analysis processes that ANSYS implements.
- As such, three ANSYS random analysis systems were derived from the first modal analysis system, one for each of the three directions of interest (along the X, Y, and Z directions), representing the two lateral and the longitudinal directions, respectively. Additionally, three ANSYS response spectrum analysis (for the shock analyses) systems were derived from the second modal analysis system, one for each direction. The first modal system was set to compute all vibrational modes falling within the frequency range of interest for the random analyses. This was based upon the nominal input loads presented in the preceding section. The same process was performed for the second modal system, to include the relevant frequency range for the shock analyses.
- Additionally, from the initial Geometry system, three new static analysis systems were derived, one for each direction, as mentioned above, within which the quasi-static analyses were defined, after having defined the material properties within it.
- Surface contacts were automatically defined in all the modules mentioned above, between all the components, a feature of the Workbench Mechanical application. Bonded contacts between all the structural components and interconnecting fasteners were implemented, to model the fact that implementing the fasteners resulted in a structural assembly with no allowable relative motion between any components.
- The structural components were discretized into finite elements through the meshing functionality. The specific type of finite element used to carry out all analyses was the 10-noded, higher order, tetrahedral element known as SOLID186, as included within the ANSYS element library [59]. This element type allows a lower number of elements to be used to discretize the structural components than what would have been necessary if lower-order elements had been used.
- The final step in the dynamic loads’ analysis methodology, across all the analysis types relevant to the current work, was to impose boundary conditions upon each model. These boundary conditions accurately reflected the physical constraints applied to the satellite while inside the launcher. As such, to model the fact that the Lightband II separation ring is connected to the launcher on its lower face, a fixed boundary condition was applied to a point defined to be in the center of the lower face of the separation ring, and which was connected to the inner edge of the lower face by rigid links. This method of support was required by the random and shock analysis systems.
3. Results
3.1. Launch Loads’ Analyses of the Perforated vs. Baseline Cases
3.1.1. Quasi-Static Loading Analysis Results
3.1.2. Random Loading Analysis Results
3.1.3. Shock Loading Analysis Results
4. Discussion and Conclusions
Author Contributions
Funding
Institutional Review Board Statement
Informed Consent Statement
Data Availability Statement
Acknowledgments
Conflicts of Interest
References
- Drenthe, N.T.; Zandbergen, B.T.C.; Curran, R.; Van Pelt, M.O. Cost Estimating of Commercial Smallsat Launch Vehicles. Acta Astronaut. 2019, 155, 160–169. [Google Scholar] [CrossRef]
- Jones, H.W. The Recent Large Reduction in Space Launch Cost. In Proceedings of the 48th International Conference on Environmental Systems, Albuquerque, NM, USA, 7–12 July 2018. [Google Scholar]
- Jones, H.W. The Future Impact of Much Lower Launch Cost. In Proceedings of the 48th International Conference on Environmental Systems, Albuquerque, NM, USA, 7–12 July 2018; pp. 1–11. [Google Scholar]
- ISO 17770: 2017; (En) Space Systems—Cube Satellites (CubeSats). International Standardization Organization: London, UK, 2017.
- Sweeting, M.N. Modern Small Satellites-Changing the Economics of Space. Proc. IEEE 2018, 106, 343–361. [Google Scholar] [CrossRef]
- Kramer, H.J.; Cracknell, A.P. An Overview of Small Satellites in Remote Sensing. Int. J. Remote Sens. 2008, 29, 4285–4337. [Google Scholar] [CrossRef]
- Xue, Y.; Li, Y.; Guang, J.; Zhang, X.; Guo, J. Small Satellite Remote Sensing and Applications—History, Current and Future. Int. J. Remote Sens. 2008, 29, 4339–4372. [Google Scholar] [CrossRef]
- Lim, J.; You, C.; Dayyani, I. Multi-Objective Topology Optimization and Structural Analysis of Periodic Spaceframe Structures. Mater. Des. 2020, 190, 108552. [Google Scholar] [CrossRef]
- Viviani, A.; Iuspa, L.; Aprovitola, A. Multi-Objective Optimization for Re-Entry Spacecraft Conceptual Design Using a Free-Form Shape Generator. Aerosp. Sci. Technol. 2017, 71, 312–324. [Google Scholar] [CrossRef]
- Cho, H.K.; Rhee, J. Vibration in a Satellite Structure with a Laminate Composite Hybrid Sandwich Panel. Compos. Struct. 2011, 93, 2566–2574. [Google Scholar] [CrossRef]
- Kuo, J.C.; Hung, H.C.; Yang, M.Y.; Chen, C.R.; Lin, J. Composite Materials Application on FORMOSAT-5 Remote Sensing Instrument Structure. Terr. Atmos. Ocean. Sci. 2017, 28, 157–165. [Google Scholar] [CrossRef]
- Du, Z.; Zhu, M.; Wang, Z.; Yang, J. Design and Application of Composite Platform with Extreme Low Thermal Deformation for Satellite. Compos. Struct. 2016, 152, 693–703. [Google Scholar] [CrossRef]
- Kwon, S.-C.; Son, J.-H.; Song, S.-C.; Park, J.-H.; Koo, K.-R.; Oh, H.-U. Innovative Mechanical Design Strategy for Actualizing 80 Kg-Class X-Band Active SAR Small Satellite of S-STEP. Aerospace 2021, 8, 149. [Google Scholar] [CrossRef]
- Dawood, S.D.S.; Harithuddin, A.S.M.; Harmin, M.Y. Modal Analysis of Conceptual Microsatellite Design Employing Perforated Structural Components for Mass Reduction. Aerospace 2022, 9, 23. [Google Scholar] [CrossRef]
- Cunningham, S.M.; Tanner, D.A.; Clifford, S.; Butan, D.; Southern, M. Effect of Perforations on Resonant Modes of Flat Circular Plates. In Key Engineering Materials; Trans Tech Publications Ltd.: Bäch, Switzerland, 2020; Volume 865 KEM, pp. 31–35. [Google Scholar]
- Abdelrahman, A.A.; Eltaher, M.A.; Kabeel, A.M.; Abdraboh, A.M.; Hendi, A.A. Free and Forced Analysis of Perforated Beams. Steel Compos. Struct. 2019, 31, 489–502. [Google Scholar] [CrossRef]
- Ghonasgi, K.; Bakal, K.; Mali, K.D. A Parametric Study on Free Vibration of Multi-Perforated Rectangular Plates. Procedia Eng. 2016, 144, 60–67. [Google Scholar] [CrossRef]
- Jeong, K.H.; Jhung, M.J. Free Vibration Analysis of Partially Perforated Circular Plates. Procedia Eng. 2017, 199, 182–187. [Google Scholar] [CrossRef]
- Formisano, A.; Lombardi, L.; Mazzolani, F.M. Perforated Metal Shear Panels as Bracing Devices of Seismic-Resistant Structures. J. Constr. Steel Res. 2016, 126, 37–49. [Google Scholar] [CrossRef]
- Sailesh, R.; Yuvaraj, L.; Pitchaimani, J.; Doddamani, M.; Mailan Chinnapandi, L.B. Acoustic Behaviour of 3D Printed Bio-Degradable Micro-Perforated Panels with Varying Perforation Cross-Sections. Appl. Acoust. 2021, 174, 107769. [Google Scholar] [CrossRef]
- Millan, R.M.; von Steiger, R.; Ariel, M.; Bartalev, S.; Borgeaud, M.; Campagnola, S.; Castillo-Rogez, J.C.; Fléron, R.; Gass, V.; Gregorio, A.; et al. Small Satellites for Space Science: A COSPAR Scientific Roadmap. Adv. Space Res. 2019, 64, 1466–1517. [Google Scholar] [CrossRef]
- CalPoly. Cubesat Design Specification; CalPoly: San Luis Obispo, CA, USA, 2009; Volume 8651. [Google Scholar]
- Jin, Y.; Shi, Y.; Yu, G.C.; Wei, G.T.; Hu, B.; Wu, L.Z. A Multifunctional Honeycomb Metastructure for Vibration Suppression. Int. J. Mech. Sci. 2020, 188, 105964. [Google Scholar] [CrossRef]
- Wagih, A.M.; Hegaze, M.M.; Kamel, M.A. FE Modeling of Satellite’s Honeycomb Sandwich Panels Using Shell Approach and Solid Approach. In Proceedings of the AIAA SPACE and Astronautics Forum and Exposition, SPACE 2017, AIAA, Orlando, FL, USA, 12–14 September 2017. [Google Scholar]
- Ontaç, S.; Daǧ, S.; Gökler, M.I. Structural Finite Element Analysis of Stiffened and Honeycomb Panels of the RASAT Satellite. In Proceedings of the 3rd International Conference on Recent Advances in Space Technologies, RAST 2007, IEEE, Istanbul, Turkey, 14–16 June 2007; pp. 171–175. [Google Scholar]
- Salem, H.; Boutchicha, D.; Boudjemai, A. Modal Analysis of the Multi-Shaped Coupled Honeycomb Structures Used in Satellites Structural Design. Int. J. Interact. Des. Manuf. 2018, 12, 955–967. [Google Scholar] [CrossRef]
- Slimane, S.; Kebdani, S.; Boudjemai, A.; Slimane, A. Effect of Position of Tension-Loaded Inserts on Honeycomb Panels Used for Space Applications. Int. J. Interact. Des. Manuf. 2018, 12, 393–408. [Google Scholar] [CrossRef]
- Wei, J.; Cao, D.; Wang, L.; Huang, H.; Huang, W. Dynamic Modeling and Simulation for Flexible Spacecraft with Flexible Jointed Solar Panels. Int. J. Mech. Sci. 2017, 130, 558–570. [Google Scholar] [CrossRef]
- Liu, L.; Wang, X.; Sun, S.; Cao, D.; Liu, X. Dynamic Characteristics of Flexible Spacecraft with Double Solar Panels Subjected to Solar Radiation. Int. J. Mech. Sci. 2019, 151, 22–32. [Google Scholar] [CrossRef]
- Rosly, N.A.; Harmin, M.Y.; Majid, D.L.A.A. Preliminary Investigation on Experimental Modal Analysis of High Aspect Ratio Rectangular Wing Model. Int. J. Eng. Technol. 2018, 7, 151–154. [Google Scholar] [CrossRef]
- Othman, M.S.; Teh, L.; Harmin, M.Y. Experimental Modal Analysis of a Simple Rectangular Wing with Varying Rib’s Orientation. In Lecture Notes in Mechanical Engineering; Springer: Singapore, 2020; pp. 473–479. [Google Scholar] [CrossRef]
- Nadkarni, I.; Bhardwaj, R.; Ninan, S.; Chippa, S.P. Experimental Modal Parameter Identification and Validation of Cantilever Beam. Mater. Today Proc. 2021, 38, 319–324. [Google Scholar] [CrossRef]
- Wertz, J.R.; Everett, D.F.; Puschell, J.J. Space Mission Engineering: The New SMAD; Microcosm Press: Hawthorne, CA USA, 2011; ISBN 1881883167. [Google Scholar]
- Sarafin, T. Spacecraft Structures and Mechanisms; 2007 (RPT); Joint Publishing: Hawthorne, CA, USA; Microcosm Press and Springer: New York, NY, USA, 1995. [Google Scholar]
- Wijker, J.J. Spacecraft Structures; Springer: Berlin/Heidelberg, Germany, 2008; ISBN 9783540755524. [Google Scholar]
- Peter, F.; Graham, S.; John, S. (Eds.) Spacecraft Systems Engineering, 4th ed.; Wiley: New York, NY, USA, 2011; ISBN 978-1-119-97836-7. [Google Scholar]
- Abdelal, G.F.; Abuelfoutouh, N.; Gad, A.H. Finite Element Analysis for Satellite Structures; Springer: London, UK, 2013. [Google Scholar]
- NASA Load Analyses of Spacecraft and Payloads. NASA Tech. Stand. Syst. 1996, NASA-STD-5, 20.
- ECSS. ECSS-E-HH-32-26A, Spacecraft Mechanical Loads Analysis Handbook; ESA Requirements and Standards Division: Noordwijk, The Netherlands, 2013; pp. 34–36. [Google Scholar]
- Aborehab, A.; Kassem, M.; Farid Nemnem, A.; Kamel, M.; Kamel, H. Configuration Design and Modeling of an Efficient Small Satellite Structure. Eng. Solid Mech. 2020, 8, 7–20. [Google Scholar] [CrossRef]
- Oh, H.-U.; Jeon, S.-H.; Kwon, S.-C. Structural Design and Analysis of 1U Standardized STEP Cube Lab for On-Orbit Verification of Fundamental Space Technologies. Int. J. Mater. Mech. Manuf. 2014, 2, 239–244. [Google Scholar] [CrossRef]
- Cote, T.; Spicer, R.; Kearns, A.; Do, N.; Soliman, H. Development and Test of an Additively Manufactured Espa Class Satellite. In Proceedings of the AIAA Scitech 2020 Forum, Orlando, FL, USA, 6–10 January 2020; Volume 1 PartF, pp. 1–16. [Google Scholar] [CrossRef]
- Okuyama, K.-I.; Hibino, S.; Matsuoka, M.; Bendoukha, S.A.; Lidtke, A. A Modification of an estimation method of the natural frequency of a cube form micro satellite. Int. J. Res. -GRANTHAALAYAH 2018, 6, 121–131. [Google Scholar] [CrossRef]
- Park, Y.-K.; Kim, G.-N.; Park, S.-Y. Novel Structure and Thermal Design and Analysis for CubeSats in Formation Flying. Aerospace 2021, 8, 150. [Google Scholar] [CrossRef]
- SpaceX Falcon 9 Launch Vehicle Payload User’s Guide; SpaceX: Hawthorne, CA, USA, 2008; p. 65.
- Arianespace Soyuz User’s Manual Issue 2 Revision 0. 2012. Available online: https://www.arianespace.com/wp-content/uploads/2015/09/Soyuz-Users-Manual-March-2012.pdf (accessed on 19 May 2022).
- ISC Kosmotras. Space Launch System Dnepr User’s Guide; ISC Kosmotras: Moscow, Russia, 2001. [Google Scholar]
- PSC. 2000785 Rev D User’s Manual for Mark II Lightband; Planetary Systems Corporation: Silver Spring, MD, USA, 2013. [Google Scholar]
- LaRocca, D.M.; Kaaret, P.; Kirchner, D.L.; Zajczyk, A.; Robison, W.; Johnson, T.E.; Jahoda, K.M.; Fuelberth, W.; Gulick, H.C.; McCurdy, R.; et al. Design and Construction of the X-Ray Instrumentation Onboard the HaloSat CubeSat. J. Astron. Telesc. Instrum. Syst. 2020, 6, 1. [Google Scholar] [CrossRef]
- Rainville, E.; Wagner, J.; Whitesel, P. Final Design Report Deployable Cover for CubeSat FUV Imager; Calif. Polytech State University: Kashiwa, Japan, 2019. [Google Scholar]
- Chan, Y.N.; Harmin, M.Y.; Othman, M.S. Parametric Study of Varying Ribs Orientation and Sweep Angle of Un-Tapered Wing Box Model. Int. J. Eng. Technol. 2018, 7, 155–159. [Google Scholar] [CrossRef]
- Tsiatas, G.C.; Charalampakis, A.E. Optimizing the Natural Frequencies of Axially Functionally Graded Beams and Arches. Compos. Struct. 2017, 160, 256–266. [Google Scholar] [CrossRef]
- Muc, A. Natural Frequencies of Rectangular Laminated Plates-Introduction to Optimal Design in Aeroelastic Problems. Aerospace 2018, 5, 95. [Google Scholar] [CrossRef]
- Othman, M.S.; Chun, O.T.; Harmin, M.Y.; Romli, F.I. Aeroelastic Effects of a Simple Rectangular Wing-Box Model with Varying Rib Orientations. IOP Conf. Ser. Mater. Sci. Eng. 2016, 152, 012009. [Google Scholar] [CrossRef]
- Dawood, S.D.S.; Harmin, M.Y.; Harithuddin, A.S.M.; Ciang, C.C.; Rafie, A.S.M. Computational Study of Mass Reduction of a Conceptual Microsatellite Structural Subassembly Utilizing Metal Perforations. J. Aeronaut. Astronaut. Aviat. 2021, 53, 57–66. [Google Scholar] [CrossRef]
- Spaceflight Inc. Spaceflight Mission Planning Guide; Spaceflight Inc.: Seattle, WA, USA, 2019. [Google Scholar]
- ECSS ECSS-E-ST-32-10C—Structural Factors of Safety for Spaceflight Hardware; The European Cooperation for Space Standardization: Noordwijk, The Netherlands, 2004; pp. 443–458.
- Beaulieu, R.A. Margin of Safety Definition and Examples Used in Safety Basis Documents and the USQ Process; NSTec LLC; Las Vegas and Mercury: Las Vegas, NV, USA, 2013. [Google Scholar]
- ANSYS. ANSYS Mechanical APDL Theory Reference; ANSYS Inc.: Washington, DC, USA, 2013; pp. 1–909, Release 15. [Google Scholar]
Factor | Value | Description |
---|---|---|
Model Factor (KM) | 1.2 | Accounts for uncertainties in mathematical models predicting dynamic response, loads, and evaluating load paths. |
Project Factor (KP) | 1.2 | Accounts for how mature a specific project is, in terms of the mass budget, the degree that the design is developed, and the level of confidence in the specifications of the project. |
Yield Factor of Safety (FOSY) | 1.25 | Accounts for the possibility of material yielding as a result of load application. |
Frequency (Hz) | Acceptance PSD (g2/Hz) |
---|---|
20 | 0.056 |
40 | 0.056 |
50 | 0.06 |
800 | 0.06 |
1300 | 0.05 |
2000 | 0.05 |
Frequency (Hz) | Acceleration (g) (11.732-inch) | Acceleration (g) (23.25-inch) | Acceleration (g) (15-inch) (Interpolated) | Acceleration (m/s2) (15-inch) |
---|---|---|---|---|
100 | 2.3 | 11 | 4.771 | 46.783 |
200 | 5 | 11.500 | 6.846 | 67.132 |
300 | 6 | 80 | 27.016 | 264.919 |
400 | 10 | 160 | 52.600 | 515.796 |
500 | 13 | 100 | 37.708 | 369.765 |
600 | 20 | 110 | 45.560 | 446.761 |
700 | 30 | 145 | 62.660 | 614.444 |
800 | 50 | 240 | 103.960 | 1019.432 |
900 | 50 | 160 | 81.240 | 796.639 |
1000 | 55 | 150 | 81.980 | 803.896 |
2000 | 180 | 210 | 188.520 | 1848.627 |
3000 | 180 | 190 | 182.840 | 1792.929 |
4000 | 200 | 205 | 201.420 | 1975.125 |
5000 | 450 | 350 | 421.600 | 4134.210 |
6000 | 600 | 480 | 565.920 | 5549.412 |
7000 | 700 | 580 | 665.920 | 6530.012 |
8000 | 500 | 700 | 556.800 | 5459.981 |
9000 | 500 | 750 | 571.000 | 5599.226 |
10,000 | 400 | 1000 | 570.400 | 5593.342 |
Baseline Case | |||
---|---|---|---|
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress | |||
2.992 × 107 N/m2 | 9.799 × 107 N/m2 | 9.860 × 107 N/m2 | |
Component with Maximum von Mises Stress | Central Box Plate +Y Facing | Base Plate | Base Plate |
Yield MoS | +6.372 | +1.251 | +1.237 |
Finalized Perforated Case | |||
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress | |||
4.069 × 107 N/m2 | 4.707 × 107 N/m2 | 4.634 × 107 N/m2 | |
Component with Maximum von Mises Stress | Side Panel +X Facing | Base Plate | Base Plate |
Yield MoS | +4.420 | +3.686 | +3.760 |
Baseline Case | ||
---|---|---|
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress | ||
2.992 × 107 N/m2 | 8.613 × 106 N/m2 | |
Component with von Mises Stress | +Y Facing Central Box Plate | +X Facing Side Panel |
Yield MoS | +6.372 | +24.608 |
Finalized Perforated Case | ||
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress | ||
6.140 × 106 N/m2 | 4.069 × 107 N/m2 | |
Component with von Mises Stress | +Y Facing Central Box Plate | +X Facing Side Panel |
Yield MoS | +34.922 | +4.420 |
Baseline Case | |||
---|---|---|---|
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress (3σ Values) | |||
8.264 × 107 N/m2 | 1.763 × 108 N/m2 | 1.765 × 108 N/m2 | |
Component with Maximum von Mises Stress | Central Box Plate −X Facing | Base Plate | Base Plate |
Yield MoS | +1.669 | +0.251 | +0.250 |
Finalized Perforated Case | |||
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress (3σ Values) | |||
2.412 × 107 N/m2 | 2.146 × 107 N/m2 | 2.142 × 107 N/m2 | |
Component with Maximum von Mises Stress | Side Panel −X Facing | Base Plate | Base Plate |
Yield MoS | +8.144 | +9.278 | +9.297 |
Baseline Case | ||
---|---|---|
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress (3σ Values) | ||
8.264 × 107 N/m2 | 1.811 × 107 N/m2 | |
Component with von Mises Stress | Central Box Plate −X Facing | Side Panel −X Facing |
Yield MoS | +1.669 | +35.529 |
Finalized Perforated Case | ||
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress (3σ Values) | ||
2.896 × 106 N/m2 | 2.412 × 107 N/m2 | |
Component with von Mises Stress | Side Panel −X Facing | Central Box Plate −X Facing |
Yield MoS | +75.160 | +8.144 |
Baseline Case | |||
---|---|---|---|
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress | |||
2.677 × 107 N/m2 | 3.217 × 107 N/m2 | 3.119 × 107 N/m2 | |
Component with Maximum von Mises Stress | Base Plate | Base Plate | Base Plate |
Yield MoS | +7.239 | +5.865 | +6.071 |
Finalized Perforated Case | |||
Longitudinal (Z-Axis) | Lateral (X-Axis) | Lateral (Y-Axis) | |
Maximum von Mises Stress | |||
2.830 × 107 N/m2 | 1.626 × 107 N/m2 | 1.619 × 107 N/m2 | |
Component with Maximum von Mises Stress | Middle Plate | Side Plate -X Facing | Side Plate +Y Facing |
Yield MoS | +6.794 | +12.565 | +12.623 |
Baseline Case | ||
---|---|---|
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress | ||
2.677 × 107 N/m2 | 1.498 × 107 N/m2 | |
Component with von Mises Stress | Base Plate | Middle Plate |
Yield MoS | +7.239 | +13.724 |
Finalized Perforated Case | ||
Longitudinal (Z-Axis) | Longitudinal (Z-Axis) | |
von Mises Stress | ||
1.982 × 107 N/m2 | 2.830 × 107 N/m2 | |
Component with von Mises Stress | Base Plate | Middle Plate |
Yield MoS | +10.128 | +6.794 |
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Dawood, S.D.S.; Harmin, M.Y. Structural Responses of a Conceptual Microsatellite Structure Incorporating Perforation Patterns to Dynamic Launch Loads. Aerospace 2022, 9, 448. https://doi.org/10.3390/aerospace9080448
Dawood SDS, Harmin MY. Structural Responses of a Conceptual Microsatellite Structure Incorporating Perforation Patterns to Dynamic Launch Loads. Aerospace. 2022; 9(8):448. https://doi.org/10.3390/aerospace9080448
Chicago/Turabian StyleDawood, Sarmad Dawood Salman, and Mohammad Yazdi Harmin. 2022. "Structural Responses of a Conceptual Microsatellite Structure Incorporating Perforation Patterns to Dynamic Launch Loads" Aerospace 9, no. 8: 448. https://doi.org/10.3390/aerospace9080448
APA StyleDawood, S. D. S., & Harmin, M. Y. (2022). Structural Responses of a Conceptual Microsatellite Structure Incorporating Perforation Patterns to Dynamic Launch Loads. Aerospace, 9(8), 448. https://doi.org/10.3390/aerospace9080448