Cbems 189 Final Report

8/13/2019 Cbems 189 Final Report

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Department of Chemical Engineering and Materials Science

Henry Samueli School of Engineering

University of California, Irvine

Senior Design Project:

Material Selection and Finite Element Analysis for

Structural Support of UCI CubeSat Spacecraft

Faculty Advisor: Farghalli A. Mohammed

Group Members: Randy Ting and Alex Yang

March 2009


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TABLE OF CONTENTS

Table of Contents iiList of Figures iii

1. Introduction

1.1 Background Information 11.2 Motivation 11.3 Purpose 1

2. Prior Work2.1 Literature Survey 2

3. Design Constraints

3.1 Material 33.2 Monetary 33.3 Technological 33.4 Manufacturability 33.5 Experience 3

3.6 Environmental Impact 43.7 Health and Safety 4

4. Design4.1 Design Processes and Consideration 44.2 Design Sketches and Drawings 5

5. Results and Discussion5.1. Materials Selection 75.2. Assumptions in Obtaining Results 75.3. Factor of Safety 95.4. Hand Calculation 105.5. Finite Element Analysis Parameters 105.6. Results 10

6. Conclusions6.1. Structural Conclusions 13

7. Plan for Future Work7.1. Destructive Testing 147.2. Proof Load Structure Schematic 147.3. Redesign Using Information from Destructive Testing 15

8. References8.1. List of Referenced Works 16

Appendix 17


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1. INTRODUCTION

The UCI Satellite team is an organization formed with one mission: to design, build, andlaunch a CubeSat spacecraft, UCISAT-1. It was formed in 2003 after professors from Stanford

and Cal Poly San Luis Obispo began a collaborative effort to create a small-scale and cost-effective platform for university space research opportunities around the world. UCIs Satelliteteam is sponsored by several different organizations including Boeing Corporation, UC IrvinesUndergraduate Research Opportunities Program, and the team members themselves. After fouryears of designing the Cubesat spacecraft, the team is approaching the fabrication phase of theproject. Currently, CAD Models of the Cubesat structure have been generated, but no analysishas been conducted on its survivability after launch. This project is an effort to design of theCubesat structure material and geometry for fabrication and implementation into the overallsystem.

1.1 Background Information

The UCI Satellite team currently has two programs that can perform finite elementanalysis: FEMAP with NX Nastran and CATIA. It also has the ability to order a test pod from

the Cubesat organization for vibration testing. Also, various machine shops are available for theteam to use for fabrication of any parts needed for the project. This project will focus mainly onthe analysis using finite element software.

1.2 Motivation

This paper provides a process overview for the finite element analysis and materialsselection process. From initial project conception to final design release, this project follows thesteps that engineers must make in everyday decisions.

FEA (Finite Element Analysis) is a tool commonly used by Engineers who are involvedwith structural design. Nearly every flavor of CAD software contains some type of FEA toolembedded within it. While some are simple, and others are complex, the results are meaninglessunless they are verified. In order to confirm that the results obtained through the computeranalysis software are true, the results obtained experimentally are compared.

1.3 PurposeThe main purpose of this project is to design the material and geometry of UCIs Cubesat

spacecraft structure using finite element software and empirical data. The end goal of the projectis to successfully fabricate the structure to be integrated into the Cubesat system to supportlaunch in the near future. The structure designed shall meet the following specifications set bythe Cubesat organization at Cal Poly San Luis Obispo:

Shall not fail under 15g load Shall not fail under vibration testing specified by Cubesat organization

Shall not weigh over 1kg Material shall not have coefficient of thermal expansion greater than AL 6061-T6 Shall not be electrically conductive

To determine the best material for the structure, the material properties of different polymers aswell as metals will be investigated. Among the important parameters being analyzed, density,strength, stiffness, and machinability will be the most critical. Finite element software will beused to simulate the environment that the structure will undergo. It will be used to predictbehavior of the complex geometry to be used as the satellites structure. Following fabrication,


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empirical testing will be performed using a compression test and a vibration test in a test podprovided by Cal Poly San Luis Obispo.

2. PRIOR WORK

2.1 Literature Survey

Article #1Jung, Park, Seo, Han, Kim: Structural Vibration Analysis of Electronic Equipment for

Satellite Under Launch Environment, Key Engineering materials, Volumes 270-273, pp1440-1445,2004

The authors investigated the result of impulse on a satellite structure on the PCB boardswhich are commonly carried on satellites. Similar to the cube structure used by the UCICubeSAT, the authors analyzed the effects of spacers on the overall effect of one PCB mountedto an aluminum frame. In performing this analysis, the authors found that connection points andmethod of connection played an important part in the modes of the FE model.

Article #2Conlon, Hambric: Predicting the vibroacoustic response of satellite equipment panels, Soc. Am,Vol. 113, No.3, pp1454-1474, March 2003

In this article, the authors investigated the effects of launch on a large scale satellite.Instead of performing an FEA analysis, the author computes and compares Statistical ElementAnalysis (SEA) methods in order to provide an alternative to the standard FEA analysis. Theauthor found that the results obtained via both methods are identical. The authors also observedthat simply lumping masses is ineffective in conveying the real behavior of a satellite structurecontaining multiple parts.Article #3Pater, Curto: Advanced Materials for Space Applications, Acta Astronautica, Vol 61, pp1121-1129, 2007

In this article, alternative materials used in space applications are discussed. Of particularsignificance to this project is PETI-5/IM7 composite. This material has been widely used indefense applications, including on the now defunct High Speed Civil Transport (HSCT) project.With a high glass transition temperature, high thermal and mechanical properties, and among thehighest tensile strengths of any composite structure, PETI-5/IM7 contains many properties whichcould complement the desired material property requirements for the UCI CubeSat.Article #4Achutuni, Menzel: Space Systems Consideration in the Design of Advanced GeostationaryOperational Enviromental Satellites, Adv. Space Res, Vol 23, No. 8, pp1377-1384, 1999

In this article, the materials of the newer age are discussed. New materials, includingGraphite/Epoxy compsites, Titanium, and Silicon Carbide have been found to be viablealternatives to classic materials, such as Aluminum and Beryllium. The major flaws with thematerials presented include out gassing in composites containing epoxy and electricalconductivity.Article #5Flint, Melcher, Hanselka, The Promise of Smart Materials for Small Satellites, ActaAstronautica, Vol 39, No 9-12, pp809-814, 1996.


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The uses of smart materials were analyzed. The alternative materials outlined within thearticle could prove to be useful in the design of the UCI CubeSat, since weight restrictions existand the potential weight savings could justify their use. Additionally, the materials outlinedwithin this paper could provide higher performance and enhance reliability of the structure.

3. DESIGN CONSTRAINTS

There are multiple types of constraints for this project; they include (but are not limited to):material, monetary, and technological, manufacturability, and experience constraints.

3.1 Material

Since a great amount of the project deals with materials which are not currently widelyused, it is expected that some materials may require significant time to acquire them.Additionally, these materials are generally scarce, and it may be difficult to use them to verifynumerical results obtained in the project analysis since they may not be available in the desired

sizes.

3.2 MonetarySince the objective is to design and build a proof load structure which will be used in

order to simulate a 15g environment, it will be necessary to purchase materials in order todevelop this proof load structure. Destructive testing can be very costly, but much engineeringanalysis will be conducted initially to ensure the data collected from destructive testing will beuseful. This structure will compromise approximately $150 in materials.

3.3 TechnologicalIn conducting an FEA, large amounts of computing power (CPU time) are required.

Similar to the compiling process in a programming language, the compilation of data from anFEA process requires a large amount of time, depending upon the amount of computing poweravailable. In some cases, geometries may not be able to be analyzed using FEA because onlydesktop computers are available rather than a workstation grade computer. For this reason, it isanticipated that computing power will be a constraint in conducting this project.

3.4 Manufacturability

Manufacturing the proof load structure will be difficult because it will assume a complex3-dimensional geometry. In order to fabricate the proof load structure, work must be outsourcedto a machinist with CNC capabilities. This will result in more time required to set up theempirical testing of the structure. The cost of outsourcing fabrication is much greater than in-house (on-campus) fabrication. The estimated cost of fabrication of the proof load structure is$500.

3.5 ExperienceWhile the fundamental concepts of vibration analysis are known to the team members, in

order to effectively perform modal analysis of the Cubesat structure, extensive research must beperformed to become accustomed to the approach for solving the vibration problems. Textbookson the modal analysis must be found so that independent learning can be accomplished.


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3.6 Environmental Impact

The fabrication of the proof load structure will produce waste in the form of excessmaterial. Depending on the material used, significant research must be conducted to ensureproper disposal of waste material. Improper disposal can result in pollution of landfills.

3.7 Health and SafetyBecause of simplicity of geometry, certain parts of the proof load structure can bemachined easily on-campus. Machining can produce particles that can be inhaled by thefabricator. In order to decrease the risk of inhaling particles produced from machining, a maskwill be worn during fabrication. Also, eye protection will be used to reduce the chance ofparticles landing in the fabricators eyes.

4. DESIGN

4.1 Design Processes and Considerations

Prior to the beginning of this project, the UCI satellite team generated a preliminaryconcept for the geometry of the Cubesat frame structure. The preliminary concept geometry wasdesigned to only support the mounting of components from other subsystems. Cuts were madein the CAD model to eliminate mass that was intuitively predicted to not be useful. A materialwas not specified for the structure. The purpose of the analysis was to determine the optimalmaterial to be used. To determine the material to be used for fabrication of the Cubesatstructure, the design process shown in Figure 4.1.1 was used.

Figure 4.1.1 Design Process Flow ChartThe CAD model of the structure was simplified to allow for easier meshing into a finite

element model. A factor of safety of 2 was chosen to mitigate any failure that could occur due tothe deviation of the prototype from the ideal finite element model. From the Cubesat


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documentation, the organizers state that a force 15 times the force of gravity should be able to beendured by the structure. From this requirement, the load was modeled as a 66.15 lb forceequally distributed between the two top surfaces of the structure assembly attached to springs.The two surfaces fastened at the bottom of the structure assembly which were fastened to springswere constrained in the finite element model to be fixed.

To determine the possible materials that would satisfy the thermal coefficient of thermalexpansion constraint set by the Cubesat organization, an Ashby chart of thermal expansioncoefficient vs. Youngs modulus was used. Any material with coefficient of thermal expansiongreater than that of aluminum was eliminated from being a possible candidate for the Cubesatstructure. A table of material properties for the possible candidates was generated to includedensity, Youngs modulus, tensile yield strength, and approximate price/weight.

Aluminum 6061-T6 was chosen as a baseline material, and a finite element analysis wasperformed using FEMAP/NX-Nastran and CATIA to ensure consistent results. The maximumsolid Von Mises stress was determined for the geometry. Materials with yield strengths less thanthat of the maximum solid Von Mises stress were eliminated as possible material candidates forthe structure. The densities of the remaining candidates were used along with the calculated

volume of the structure to eliminate any materials which would result in a structure that exceedsthe maximum 1 kg mass constraint set by the Cubesat organization. Youngs moduli of theremaining materials were compared to determine the stiffest material which would result inminimum deflection of holes. This ensures that load will not be transferred to any componentsthat are not part of the structures subsystem.

4.2 Design Sketches and Drawings

The design sketches and drawings for the UCISAT satellite can be found in Appendix A.

5. RESULTS AND DISCUSSION

5.1 Materials SelectionTo narrow down the search for feasible materials for use in the Cubesat structure, an

Ashby materials selection chart was used. The maximum thermal expansion coefficientconstraint set by the Cubesat organization was used first to filter the vast selection of materials.


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Figure 5.1.1 Ashby Chart Used to Filter High Thermal Expansion Coefficient Materials

An Ashby Chart was used to filter materials with high coefficients of expansion. A line wasdrawn along the constant thermal expansion coefficient of Aluminum alloys.

The elimination process for the materials was as follows: All materials above the line were rejected as feasible solutions because of high thermal

expansion coefficients.

The families of materials that remained as feasible solutions were natural materials,ceramics, metals, and composites.

The fabrication capability of ceramics was researched, and it was concluded thatceramics should be eliminated as candidates because it is difficult to machine. It was alsofound that the ceramics are a brittle material, which would increase its chances of failingonce its yield strength is reached.

Stone and brick were found to not be a feasible solution because their material propertiescould not be predicted or controlled easily.

Carbon fiber reinforced plastics were also eliminated as feasible solutions becauseresearch results showed that their mechanical properties were anisotropic.

The project team lacked the resources to analyze the behavior of complex three-dimensional geometries composed of anisotropic materials under load.

The fabrication cost of CFRPs was determined to be greater than that of the projectsbudget.

Metals were kept as feasible solutions because they are capable of good machinability,easy to obtain, and generally isotropic.


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Table 5.1.1 Typical Metals Mechanical Properties Chart

Material Properties of typical, easy to obtain metals were researched to be down selected.The mechanical properties of the metals are tabulated in Table 5.1.1. To determine if eachmaterial was capable of satisfying the 1kg mass constraint, the approximate mass for eachmaterial was calculated using a volume of 7.21 cubic inches as determined by the SolidWorksCAD model generated. The 7.21 cubic inches volume was only considering the volume of theframe structure components without any electrical components. Copper Alloy C17000, Cast Iron

60-40-18, and 1018 steel were determined to result in a frame structure which was too heavy toallow a weight margin for accommodating electrical components. Therefore, 6061 AluminumT6 and Titanium Alloy ASTM Grade 1 were considered for final analysis in FEA. Aluminumand titanium can be anodized to prevent the conduction of electricity that could short circuitelectronics components. Empirical tests were not performed to determine whether or not ananodized surface would be sufficient to prevent electrical conduction, but conductivity values

were reduced by a factor of 103.

5.2. Assumptions in Obtaining ResultsIn order to more effectively analyze the structure of the satellite, assumptions needed to

be made which simplified the analysis.

Since the side panels of the CUBESAT satellite are used as a structural element, theconnection between these points and the support frames needed to be analyzed. The modelemploys button-head cap screws in conjunction with countersunk screws in order to affix thepanels to the structure. Button-head cap screws hold the side, back, and front panels to thesupport frames. An illustration can be found on the following page on Figure 5.2.1.

Material Elastic Modulus

(ksi)

Tensile Yield

Strength (ksi)

Density

(lbm/in^3)

Average Thermal

Expansion

Coefficient (in/in F)

Mass

(kg)

6061 A luminum T6 1.00E+04 35 0.098 13.1 0.32

Copper Alloy C17000 1.90E+04 37 0.304 9.4 1.00

Cast Iron 60-40-18 2.45E+04 47.7 0.246 5.55 0.81

1018 Steel 2.90E+04 32 0.28 4.8 0.92

Titanium Alloy ASTM Grade 1 1.49E+04 30 0.163 5.1 0.53


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Figure 5.2.1. Button-Head Screw Attachments

The illustration above shows the typical arrangement of button-head cap screws on the UCISATmodel. There are (5) instances of the set of (4) throughout the entire structure. All of the panels,with the exception of the top panel, employ this fixation method. In conducting the analysis, arigid surface connection was assumed for each of the connectors. The rigid surface connectiontook the shape of a circle, with the same diameter as the flanged surface of the button-headscrew. The connection is an interaction between the plate and the support frame.

In addition to button-head cap screws, countersunk screws are also employed on the toppanel of the UCISAT structure. They are illustrated on Figure 5.2.2 below.

Figure 5.2.2. Countersunk Screw Locations

The countersunk screw locations were approximated in the model using a projection of theflanged face upon the mating surface of the panel and the support structure. There are (8)


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instances of this connection on our top plate. A rigid surface connection with a circular areaequal to the projected area was assumed in order to complete the analysis.

Assumptions regarding the forces were also considered during this analysis. It wasassumed that a 15g load was applied to the spring plunger points, which are shown in Figure5.2.3.

Figure 5.2.3. Spring Plunger Location Points.

It was assumed that the plungers would remain in contact with the launch mechanism throughoutthe duration of the launch. For this reason, the 15g force was split among the two plungers. Itwas also assumed that the opposing surface was held in a fixed position for the analysis.

5.3. Factor of Safety

A factor of safety was applied due to uncertainties in the conduction of the analysis. Thederivation of the factor of safety can be found below on Figure 5.3.1.

Contribution Value Comments

Material 1.1 The material purchased to manufacture the panels is not certified with aheat number. For this reason, only the expected properties from thedatasheets can be used.

Load Stress 1.2 Both members of the group have limited experience in stress analysismethods.

Geometry 1.0 The panels will be manufactured using 5-Axis CNC Machining (frames)and WaterJet (panels). These processes can hold true tolerances of lessthan .003

FailureAnalysis

1.1 The stresses were derived from the free body diagram, which states the

Reliability 1.4 Since only one satellite will be launched, the reliability is very important.

Figure 5.3.1: Chart Showing the Selection for Factor of Safety.


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It was found that the factor of safety was required to be 2.03. The derivation for the factor ofsafety can be found in Appendix B.

5.4. Hand Calculation

Figure 5.3.1. Hand Calculation for Stresses

As stated in the previous section, the forces were assumed to only exist on the two springplungers which were attached to the two support structures. The figure above shows a crosssection view of only one of the panels. In order to better approximate the amount of force whichis endured by the internal cross sectional area, the force was modeled and the stresses werecalculated. Since the values for the stresses were less than the yield strengths of both materialsbeing considered, we continued with the FEA analysis of the two structures together.

5.5. Finite Element Analysis Parameters

During the FEA, the OCTOTREE mesh utility was used on all of the geometricalelements of the model. The mesh size was set to .050, or 50 thousandths of an inch. A uniformforce of the following magnitude was applied.

= (. . ) = (2.03)(1 ) 9.8 = 19.894

This stress was distributed evenly upon both of the plunger mounts during the analysis.


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Additionally, a fixed plane was needed to perform the analysis. For our study, we chosethe plane which sat parallel to the plane normal to the tip of the vector which models thedirection of the force.

5.6. Results

The above parameters were applied, and the structural analysis was performed. Sketches whichdetail the principal stresses, deflection, and nodes of the frames can be found on the followingpages.

Figure 5.5.1. Principal Stresses for Structural Frame

On the plot above, the stresses in units of pounds per square inch are conveyed. The meantensile stress of the structure is -38.8 psi. The maximum stresses in compression and tension are -225 psi and 148 psi, respectively.


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Figure 5.5.2. Displacement Depiction for Nodes.

In the figure above, the displacement of nodes (shown in figure 5.5.3) is shown above.From the above figure, the bottom third of the box closest to the fixed plane (shown in blue) ispredicted to have zero displacement. The maximum displacement of the nodes was found to be.214 thousandths of an inch. This maximum is located on the surface where the load is applied,and is shown in red.


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Figure 5.5.3. Node Plot for FEA Model

The figure above shows the nodes which were used to conduct the finite elementanalysis. They show the location of the finite elements analyzed throughout the structure.

6. CONCLUSIONS6.1 Structural Conclusions

After investigating the structure, and using the FEA analysis tools, we found that themaximum A table of the key findings can be found below on Figure 6.1.1.

Table 6.1.1: Maximum Values for Deflection and Stress

Parameter Value

Maximum Deflection 2.14 104Maximum Stress 148

The above table shows the maximum deflection and maximum stress which areexperienced by the UCISAT structure.

Since all of the applicable constraints for the structure were met, we concluded that 6061-T6 wasan acceptable material for use for the stress analysis.


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7. PLAN FOR FUTURE WORK

7.1. Destructive Testing

A proof load structure will be contstructed and used to determine the maximum amountload which can be applied to the UCISAT fixture. The actual loads which cause the UCISAT

structure to fail can then be determined. With this information, a designer will re-design the boxin order to meet the minimum requirement to minimize weight, producing the optimal boxpossible.

7.2. Schematic for Proof Load Structure

The purpose of a proof load structure is to accurately reproduce the forces which exist onthe structure. As modeled in Figure 5.2.3, the forces on the structure are evenly distributed on thetwo spring plungers. A graphical rendering of a proof load structure can be found below onFigure 7.3.1.

Figure 7.3.1: Proof Load Structure Graphical Representation

The graphical rendering above is the proof load structure which will be used to simulateloads on the UCISAT structure. The satellite structure will be sandwiched between the top andbottom plates, which are free to slide in the Z-axis. The UCISAT structure will be placedbetween the top and bottom plates of the proof load structure, and mass will be added to the topplate until failure in the UCISAT structure is achieved.

The ultimate tensile strength of the proof load structure will be found and calculated. Thisdata will be collected for the re-design process which is outlined in Section 7.3.


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7.3. Redesign Using Information from Destructive Testing

After destructive testing is complete, the failure mechanism will be examined. Plasticdeformation of the box is expected, with a brittle fracture failure mechanism. The direction of thecrack propagation will be carefully examined and analyzed. Material in areas where cracks areminimal will be removed from the panels in order to eliminate mass from the structure. After the

panels have been re-designed, the model will be re-analyzed using the finite element analysismethod in order to ensure that all of the constraints outlined by the project definition are stillbeing met, and the box will be produced.


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8. REFERENCES

8.1. List of Referenced Works1. Jung, Park, Seo, Han, Kim: Structural Vibration Analysis of Electronic Equipment for

Satellite Under Launch Environment, Key Engineering materials, Volumes 270-273,

pp1440-1445,20042. Conlon, Hambric: Predicting the vibroacoustic response of satellite equipment panels,Soc. Am, Vol. 113, No.3, pp1454-1474, March 2003

3. Pater, Curto: Advanced Materials for Space Applications, Acta Astronautica, Vol 61,pp1121-1129, 2007

4. Achutuni, Menzel: Space Systems Consideration in the Design of AdvancedGeostationary Operational Enviromental Satellites, Adv. Space Res, Vol 23, No. 8,pp1377-1384, 1999

5. Flint, Melcher, Hanselka, The Promise of Smart Materials for Small Satellites, ActaAstronautica, Vol 39, No 9-12, pp809-814, 1996.

6. Ashby, Michael. Materials Selection in Mechanical Design. Massachusetts, PergamonPress, 1992.7. Callister, William D. Materials Science and Engineering: an Introduction. NewYork,Wiley, 2006.


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Appendix A : UCISAT Sketches


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4X .1064X .063

42X .250

0

.350

.253

.153

.603

1.053

1.503

1.953

2.403

2.853

3.303

3.555

3.653

3.753

3.905

0

.350

.453

.953

.828

1.2

78

1.7

28

2.1

78

2.6

28

3.0

78

2.9

53

3.4

53

.202

3.7

03

3.9

05

3.5

55

3.905

.063

12345678

8 7 6 5 4 3 2 1

THE INFORMATION CONTAINED IN THIS

DRAWING IS THE SOLE PROPERTY OF

UCISAT. ANY REPRODUCTION IN PART

OR AS A WHOLE WITHOUT THE WRITTEN

PERMISSION OF UCISAT IS PROHIBITED.

PROPRIETARY AND CONFIDENTIAL

DIMENSIONS ARE IN INCHES

TOLERANCES:

FRACTIONAL .01

ANGULAR: MACH BEND

TWO PLACE DECIMAL .01

THREE PLACE DECIMAL .005

INTERPRET GEOMETRIC

TOLERANCING PER:

MATERIAL

FINISH

Aluminum 6061

COMMENTS:

UCISATTITLE:

SIZE

BPROJECT

WEIGHT:SCALE: 1:2

UNLESS OTHERWISE SPECIFIED:

A

Bottom Panel

SHEET 1

UCISAT-1DO NOT SCALE DRAWING


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R.063

R.050

R.063

R.063

R.050

R.063

R.063

R.063

R.063

R.063

R.063

R.063

R.031 R.031

R.040

R.040

R.040

R.040

2-56 UNC THRU ALL

8X .070 THRU ALL

0

.335

.469

.600

.679

.739

1.

939

1.

999

3.

199

3.

259

3.

337

3.

469

3.

602

3.

937

0

.344

.484

.584

.829

.939

3.679

3.529

3.884

3.984

4.469

.539

3.929

4.124

.255 X 82

2X .089 THRU

0-80 UNF .120

2X .047 .160

2-56 UNC .172

4X .070 .250

R.063

R.063

R.125R.125

.585

2-56 UNC .172

2X .070 .226R.040

R.040

R.063

R.063

R.063

R.031

0

.469

1.

044

.969

.679

.739

1.

939

1.

999

2.

894

2.

969

3.

199

3.

259

3.

469

R.063

R.040

R.063

R.063

R.031

R.031

R.063

R.031

R.031

0

.

335

.

469

.

462

.

669

.

585

.

835

.

844

.094

12345678

8 7 6 5 4 3 2 1



UCISAT. A NY REPRODUCTION IN PART





TOLERANCES:

FRACTIONAL .01

ANGULAR: MACH BEND



INTERPRET GEOMETRIC

TOLERANCING PER:

MATERIAL

FINISH

Al 7075-T73

COMMENTS:

UCISATTITLE:

SIZE

BPROJECT

WEIGHT:SCALE: 1:2


A

Left Frame

SHEET 1



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R.040 R.040

R.063

R.063

R.063 R.063

R.063

R.063

R.125

R.063

R.063

R.063

R.050

R.063

R.050

R.063

2-56 UNC THRU ALL8X .070 THRU ALL

R.031

0

.335

.469

.679

.739

1.

939

1.

999

3.

199

3.

259

3.

469

3.

602

3.

937

0

.344

.484

.584

.709

1.439

.539

3.559

3.884

3.929

4.124

4.469

3.429

3.709

R.063

.255 X 822X .089 THRU

2-56 UNC .170

4X .070 .250

R.040

R.040

R.063

R.031

0

.534

1.141

1.234

2.034

2.434

2.984

3.234

3.327

3.934

4.124

4.469

3.929

3.059

1.459

.344

2.609

1.859

2-56 UNC .170.070 .260

0

.167

.205

.469

.679

.739

.585

.919

1.

169

1.

939

1.

999

2.

769

3.

019

2.

894

1.

044

3.

199

3.

259

3.

469

3.

937

.080

3.

857

0 .469

.969

.679

.739

1.

044

1.

939

1.

999

2.

969

.919

1.

169

2.

769

3.

019

3.

199

3.

259

3.

469

3.

937

12345678

8 7 6 5 4 3 2 1


DRAWING IS THE SOLE PROPERTY OFUCISAT. A NY REPRODUCTION IN PARTOR AS A WHOLE WITHOUT THE WRITTEN



DIMENSIONS ARE IN INCHESTOLERANCES:

FRACTIONAL .01

ANGULAR: MACH BEND

TWO PLACE DECIMAL .01THREE PLACE DECIMAL .005

INTERPRET GEOMETRICTOLERANCING PER:

MATERIAL

FINISH

Al 7075-T73

COMMENTS:

UCISATTITLE:

SIZE

BPROJECT

WEIGHT:SCALE: 1:2


A

Right Frame

SHEET 1



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4X .106

2X R.063

12X R.090

.300

0

.240

.440

1.045

1.459

1.890

2.735

2.321

3.340

3.540

3.780

0

.100

.440

.674

1.2

08

1.6

00

2.5

26

2.7

60

1.9

92

3.1

00

3.2

00

3.0

75

.125

3.780

.063

12345678

8 7 6 5 4 3 2 1



UCISAT. ANY REPRODUCTION IN PART





TOLERANCES:

FRACTIONAL .01

ANGULAR: MACH BEND



INTERPRET GEOMETRIC

TOLERANCING PER:

MATERIAL

FINISH

Aluminum 6061

COMMENTS:

UCISATTITLE:

SIZE

BPROJECT

WEIGHT:SCALE: 1:2


A

Side Panel

SHEET 1



25/29

4X .106

2X R.063

42X .250

0

.202

.453

.828

1.

278

1.

728

2.

178

2.

628

3.

078

3.

453

3.

555

.350

3.

905

3.

703

0

.350

.253

.603

1.053

1.503

1.953

2.403

2.853

3.303

3.653

3.905

3.555

3.905

.063

12345678

8 7 6 5 4 3 2 1

THE INFORMATION CONTAINED IN THISDRAWING IS THE SOLE PROPERTY OF

UCISAT. ANY REPRODUCTION IN PARTOR AS A WHOLE WITHOUT THE WRITTEN



DIMENSIONS ARE IN INCHESTOLERANCES:

FRACTIONAL .01

ANGULAR: MACH BENDTWO PLACE DECIMAL .01THREE PLACE DECIMAL .005

INTERPRET GEOMETRICTOLERANCING PER:

MATERIAL

FINISH

Aluminum 6061

COMMENTS:

UCISATTITLE:

SIZE

BPROJECT

WEIGHT:SCALE: 1:2


A

Top Panel

SHEET 1



26/29


27/29

AppendixB:FactorofSafety


28/29


29/29

Date post:	04-Jun-2018
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Cbems 189 Final Report

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