phase 1 for spacecraft design

7/25/2019 phase 1 for spacecraft design

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PHASE 1MANUFACTURING FEASIBILITY PHASE WITH PRIMARY

FOCUS ON OPTIMIZATION OF THE ULTRA-LIGHTWEIGHTCORE MATERIAL FABRICATION PROCESS AND PRODUCTION

OF SAMPLES FOR MECHANICAL PROPERTY ASSESSMENT BY

NASA.

BY

MAHESHWOR KC

SAI KUMAR BOMULE

SAI VEERENDRA POTLA

RAKESH BADRAJISYED ALI


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TARGET CORE PROPERTIES

Property Targeted core

material

properties

D en si ty , l b/ f t3

L es s t ha n 3

C el l S iz e, i n N /A

Compressive

Strength, psi

1000

Compressive

modulus, ksi

295

C r us h S t re n gt h , p s i 1 0 00

Plate shear

s tr en gt h ( L) , p si

640

Plate shear

m od ul us ( L) , k si

102

Plate shearstrength (W), psi

370

Plate shear

modulus (W), ksi

38

Co st of 4 X 8 X 1

panel


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INTRODUCTION

Increasing the core thickness

greatly increases the stiffness of

the honeycomb construction,

while the weight increase is

minimal.


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Flex core

Formability feasibility with reduced anticlastic

curvature and without buckling the cell walls.

Curvatures of very tight radii are easily formed.

Higher shear strength than hexagonal

honeycomb (when formed in tight radii).

HONEYCOMB CELL CONFIGURATION


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Aramid Fiber Honeycomb:

Use HRH-10 aramid fiber for core material.

dipped in a heat-resistant phenolic resin

Features:

High strength and toughness in a small cell size

Low density nonmetallic core.

CORE MATERIAL


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Table : Mechanical properties of HRH-10

MECHANICAL PROPERTIES OF HRH-10


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FABRICATION OF HONEYCOMB CORE

(STRENGTH OPTIMIZATION OF HRH-10)

Where,

Ef = the elasticity modulus of the skins

Ec = the in-plane elasticity modulus of the core

Tf = the thickness of the skins

Tc= the thickness of the core

B= the width of the beamd = is the distance between the centers of the two faces

e = the position of the neutral axis.

Figure : Typical cross section of

Honeycomb Sandwich


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Assumption: weak core and strong CFRP skin and using classic Bernoulli hypothesis

The position of the neutral axis

Flexural Rigidity


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MECHANICAL PROPERTIES AND TEST METHODS

1. Compressive strength test

Bare compressive test-without skin to core bonding

Stabilized compressive test- with skin to core bonding


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CRUSH STRENGTH

When the spacecraft using such sandwich panels is hit by hyper velocity

meteoroids the time while the panel going plastic deformation before

getting crushed will be crucial for the astronauts to get out of the space

crafts.


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2. Plate shear test method

The Shear strength and modulus values

dimension as 6" x 2" x 0.5" for non-metallic honeycomb fibers.

initiated by bonding the specimens to 1/2" thick steel loading plates.

loading rate is normally 0.020 inches per minute.

Shear deflections are to be measured with a displacement

transducer that senses the relative movement of the two plates.

Shear modulus is calculated from the slope of the initial

straight-line portion of the load-deflection curve.

Beam-flexure test can be used to test the core with higher

densities but our goal is used to design core with density

less than 3 pcf so we ignore this test.


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SELECTION OF SKIN

We made the decision to use IM10 12K epoxy composite with continuous

unidirectional fiber

Epoxy IM10 12K composite had high strength-to-weight and high stiffness-

to-weight properties and is stronger than steel, lighter than aluminum and

as stiff as titanium.


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Unidirectional prepreg tapes have been mostly used in the aerospace industry.

Fibers that have been oriented in 0/90 degree pattern to double the strength as

compared to unidirectional orientation.

use of only one layer of prepregs have higher chance of dimpling.

ORIENTATION OF FIBERS


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UNIDIRECTIONAL CFRP PREPREG TAPE


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The table above shows the tensile and compressive properties of IM10 12k at different orientations.

SKIN ORIENTATIONS MECHANICAL PROPERTIES


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REQUIREMENTS OF ADHESIVES

adhesives must function in adverse range of temperatures.

should resist the radiations and thus the micro cracking.

Insulation, high de-bonding energy.

Adhesive should pass the outgassing standard test

procedure according to ASTM E595.


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ASTM TEST FOR OUTGASSING


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SUMMARIZED RESULTS OF ASTM E595

If (Collected volatile condensable material)CVCM 0.10% and TML

(Total mass lost) 1.0%, the material passes.

If CVCM 0.10% and TML > 1.0%, the material can pass if TML WVR

1.0%.

If CVCM > 0.10%, the material fails.

If TML WVR (Water vapor regained) > 1.0%, the material fails.


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TWO PART EPOXIES

offer excellent cohesion, resist chemicals, bonds very well with most

materials

operate over adverse range of temperature (4K) to 550F.

meet or exceed NASAs outgassing specification.

~40 g/m2 adhesive per skin considered to optimize the strength-to-weight

ratio in bending of sandwich structure.


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ADHESIVE FILLETS


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CALCULATIONS

Proposed Thickness of the core (tc)= 0.70

Proposed Weight of the core(Wc): Density*Volume = 1.8 ft3*(0.7/12*10X11)= 11.55lb

For skin (IM 10, 12k)

Area covered by 2lb spool of IM 10=11*2800m=8420.8 ft2

Weight of 11X10 IM 10 for one face of the panel= 2*110/8420.8=0.026lbs

We propose to use 8 skins layer on each face of the core and 40g/m2 (0.008 lb/ ft2

adhesive per skin layer to get the optimal strength-to-weight ratio in bending of

sandwich structure.

Total weight of the two part epoxy used in the panel= 0.008*8 (gaps between

layers)*2(faces)*110(nominal area for the panel)=14.08 lbs

Total weight of the proposed honeycomb core sandwich panel= 0.026*16(total no.of layers)+14.08+11.55=26.046lbs


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CALCULATION CONTD

Density of the overall proposed sandwich panel= 26.046/9.167=2.84lb/ ft3

Overall density < 3lb/ ft3 (OK)

Flexural rigidity

From the table 2,

Diameter of the each tow (layer of skin)= Sqrt(0.18*4/3.14)=0.478mm=0.0188

Total thickness of skin on each side of the core (8 on each side)=8*0.0188=0.1504

Total depth of the proposed panel: (0.1504)*2+ 0.70= 1.0004~1

From Eq. 1

Assuming Ec


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CONCLUSIONS


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phase 1 for spacecraft design

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