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祝飛鴻

衛星結構設計

5/31/2007

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1. What are key constraints for the spacecraft structure design?

2. How the structure design is affected by other subsystems?

3. How the structure design affects the performance of other subsystems?

4. How to distinguish a good and bad spacecraft structure design?

Pre-Class Assignment

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Spacecraft Structure Design:• What are the main functions?

• What factors need to be satisfied?

• What are major tasks?

• How to verify the design?

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Structure subsystem holds all other subsystems together:

Carry Loads - provide support all other subsystems and attach the spacecraft to launch vehicle.

Maintain geometry – alignment, thermal stability, mass center, etc.

Provide radiation shielding

The first Taiwan designed satellite

Structure design is affected by all the other subsystems

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Spacecraft structure design has to satisfy the following factors:

1. Size

2. Weight

3. Field-of-view

4. Interference

5. Alignment

6. Loads

The first Taiwan designed satellite

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衛星尺寸限制 :•Falcon 1 (Dia. 1371)•Falcon 1E (Dia. 1550)•Taurus-63 (Dia. 1405)

Falcon 1

Falcon 1E

Taurus-63

1371

1405

1550

1. Size: Fit into the fairing of candidate

launch vehicle. Provide adequate space for

component mounting.

13mm clearance

11mm clearance

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2. Weight: Not to exceed lift-off weight of the selected launch vehicle to the

desired orbit. Trade will be performed to determine the launch vehicle injection

orbit for best weight saving.

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3. Field-of-view (FOV): Define by other subsystems, e.g. attitude control

sensors, payload instruments, antenna subsystem, etc.

X Band Antenna FOV

110 °65 °65 °110 °

MSI FOV= 6 °

Star Camera FOV= 6.7° on short axis

9.2° on long axis

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4. Interference: With the launch vehicle fairing. Between components for physical contact

and assembly. Falcon-1Envelope

Solar Panel19mm clearance

X-Band Ant15.5mm clearance

GPS Ant.8.6mm clearance

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5. Alignment: Define by other subsystems, e.g. attitude control sensors,

payload instrument, etc. On ground alignment, if necessary. On-orbit thermal & hydroscopic distortion.

Requirement

Star Camera

Orientation

± 0.5

Thruster Orientation ±1.5

X-antenna Orientation ±5

S-antenna Orientation ±5

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6. Loads: Environmental loads for structure design. Loads for components and payloads.

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The structure design may not be able to satisfy all the design factors.

Therefore

Factors to be satisfied for structure design is not

a one way street

Factorsto be

satisfied

StructureDesign

System Performance

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Major tasks for spacecraft structure design include:

1. Configuration design

2. Material Selection

3. Environmental loads

4. Structure analysis

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1. Configuration Design: To accommodate all the components in a limited space while satisfying its functional requirements, every spacecraft will end up with a unique configuration.

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The First Taiwan Designed Satellite

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2. Material Selection:

Factors to be considered:

Strength-to-weight ratio

Durability

Thermal stability

Thermal conductivity

Outgassing

Cost

Lead time

Manufacture

Commonly used material:

Metals – Aluminum, etc.

Composites

Ceramics

Polymers

Semiconductors

Adhesives

Lubricants

Paints

Coating

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3. Environmental Loads: To successfully deliver the spacecraft into the orbit, the launcher has to go through several stages of state changes from lift-off to separation. Each stage is called a “flight event” and those events critical to the spacecraft design is called “critical flight events”.

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3. Environmental Loads: Each flight event will introduce loads into the spacecraft. Major types

of loads include: Transient dynamic loads caused by the changes of acceleration state

of the launcher, i.e. F = ma. F will be generated if a or m is

introduced. Random vibration loads caused by the launcher engine and aero-induced

vibration transmitted through the spacecraft mechanical interface. Acoustic loads generated from noise in the fairing of the launcher, e.g.

at lift-off and during transonic flight. Shock loads induced from the separation device.

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3. Environmental Loads: The above mentioned launcher induced loads are typically defined in

the launch vehicle user’s manual. However, these loads are specified

at the spacecraft interface except for acoustic environment. The loads

to be used for the spacecraft structure design has to be derived.

For picosat design, if P-POD is used, please refer to “The P-POD

Payload Planner’s Guide” Revision C – June 5, 2000 for definition of

launch loads.

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Environmental Loads: Among all the launch loads, the derivation of transient dynamic

loads is most involved and typically is the dominate load for

spacecraft primary structure design. Unfortunately the transient dynamic loads are structure design

dependant, e.g. magnitude of loads depends on the spacecraft

structure design (see appendix for explanation). However, loads

are required for the design. Typically spacecraft structure are designed with the quasi-static

load factors defined in the launch vehicle user’s manual, e.g. 2g

lateral and 7g axial. These quasi-static loads are only applicable if the stiffness design

of the spacecraft is above the minimum frequency requirement as

specified in the launch vehicle user’s manual, e.g. >20Hz lateral.

These loads may not be applicable for light weight second appendages,

e.g. solar panel, antenna, etc. and needs to be verified by the coupled

loads analysis.

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Coupled Loads Analysis: The natural frequencies of a spacecraft can be predicted by mathematical

model, e.g. finite element model. This model will be delivered to the

launcher supplier for coupling with the launch vehicle model. Dynamic

analysis can be performed using this combined model and critical responses

of the spacecraft can be derived for the spacecraft structure design.

Spacecraft Model

Launch VehicleModel

CombinedModel

DynamicAnalysis

Forcing Functionsof

Critical Flight Events

SpacecraftResponses

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Typical CLA Results

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Dynamic Coupling

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Structure Analysis

4. Structure Analysis:4.1 Mass property analysis

4.2 Structure member and load path

4.3 Dynamic and stress analysis

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4.1 Mass Property Analysis: One of the important factors associated with the mechanical layout

is the mass property analysis, i.e. weight and moment of inertia

(MOI) of the spacecraft. Mass property of a spacecraft can be calculated

based on the mass property of each individual

elements e.g. components, structure, hardness,

etc. The main purpose of mass property analysis

is to assure the design satisfies the weight

and CG offset constraints from the selected

launcher.

W1

W2 X

Y

D2

D1

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0 200 400 600 800 1000 1200 1400

Spacecraft Weight (lb)

2.5

2.0

1.5

1.0

0.5

0.0

Lateral CG centerline offset (in)

Falcon-1 Launcher

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4.2 Structure Members and Load Path: The spacecraft is supported by the launcher interface therefore

all the loads acting on the spacecraft has to properly transmitted

through the internal structure elements to the interface. This load

path needs to be checked before spending extensive time on

structural analysis.

No matter how complex the structure is, it is always made of basic

elements, i.e. bar, beam, plate, shell, etc.

Components => Supporting Plate => Beam => Supporting Points

Plate

Beam

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4.3 Dynamic & Stress Analysis: Finite element analysis is the most popular and accurate method to

determine the natural frequencies and internal member stresses of a spacecraft. This analysis requires construction of a finite element model.

Once the environmental loads, configuration and mass distribution have been determined, analysis can be performed to determine sizing of the structure members. Major analysis required for spacecraft structure design include dynamic (stiffness) and stress (strength) analysis.

Major goal of the dynamic analysis is to

determine natural frequencies of the

spacecraft in order to avoid dynamic

coupling between the structure

elements and with the launch vehicle.

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Dynamic & Stress Analysis: Purpose of the stress analysis is to determine the Margin of Safety (M. S.) of structure elements:

Allowable Stress or Loads M. S. = - 1 0

Max. Stress or Loads x Factor of Safety

Allowable stresses or loads depends on the material used and can be obtained from handbooks, calculations, or test data. Maximum stress or loads can be derived from the structure analysis. Factor of Safety is a factor to cover uncertainty of the analysis. Typically 1.25 is used for yield stress and 1.4 for ultimate stress.

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4.3 Dynamic & Stress Analysis: Construction finite element model of a spacecraft is a time consuming

task. Local models, e.g. panel and beam models, can be used to

determine a first approximation sizing of the structure members.

close form solution(Simply supported platewith uniform loading)

Finite element solution(Simply supported platewith concentrated mass)

close form solution(beam with concentrated force)

reaction force

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Structure design is an iterative process

However

Major design changes will have significant impact to the program

SDR(System Design Review)

PDR(Preliminary Design Review)

CDR(Critical Design Review)

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How to verify spacecraft structure design?

Mechanical Layout – Assembly and integration

Alignment – Alignment measurement

Mass Property – Mass property measurement

Quasi-static Loads – Static load test

Transient Dynamic Loads – Sine vibration test

Random Vibration Loads – Random vibration test

Acoustic Loads – Acoustic test

Shock Loads – Shock test

On-orbit loads – Thermal vacuum test

Depends on the program constraints and risk assessmentnot all the tests are required.

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Homework Problem

1. Revise answer to the pre-assignment problems.

2. Define detailed step by step process for your picosat

structure design. Identify sources for the required

inputs.

Please provide your answer by 6/8 (Fri)

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What you have learned is:

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Reference

Spacecraft Systems Engineering, 2nd edition, Chapter 9,

Edited by Peter Fortescue and John Stark, Wiley

Publishers, 1995.

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Appendix

Phenomena of Dynamic Coupling

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Dynamic Coupling

Among all the launch loads, the derivation of transient

dynamic loads is most involved and typically is the

dominate load for spacecraft primary structure design.

To understand the derivation of transient dynamic loads,

the concept of “dynamic coupling” needs to be explained.

Based on the basic vibration theory, the natural frequency

of a mass spring system can be expressed as:

1 f = ------ K/M 2

Where

f = natural frequency (Hz: cycle/second)

M = mass of the system

K = spring constant of the system

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Dynamic Coupling Based on the above equation, a spring-mass system with K1 = 654,000 lb/in and weight W1= 4,000 lbs will have f1 = 40Hz (verify it!). Assume a second system has f2 = 75Hz. (if this system has 30 lbs weight, what should be the value of K2?) The forced response of these two systems subjected to 1g sinusoidal force base excitation with 3% damping ratio will have 16.7g response at their natural frequency, i.e. For system 1: 16.7g at 40Hz For system 2: 16.7g at 75Hz

(Please refer to any vibration text book for derivation of results)

W

K

1g

a

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Dynamic Coupling

Suppose we stack these two system together, the response

of the system can be derived as:

39.8Hz 75.4Hz a1 16.6g 0.4g

a2 23.1g 6.4g

where 39.8Hz and 75.4Hz are the natural

frequencies of the combined system. (Please refer to advanced vibration text book

for derivation of results)

W2

W1

K2

K1

1g

a1

a2

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Dynamic Coupling

Now, let’s change the second system to have natural

frequency of 40Hz, then the responses will be:

38.3Hz 41.8Hz a1 9.9g 9.2g

a2 99.2g 83.4g

where 38.3Hz and 41.8Hz are the natural

frequencies of the combined system.

W2

W1

K2

K1

1g

a1

a2

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Dynamic Coupling

It can be seen that by changing the natural frequency

of the second system to be identical to the first

system, the maximum response of the second

system will increase from 23.2g to 99.2g.

This phenomenon is called “dynamic

coupling”. The more closer natural

frequencies of the two systems, the

higher response the system will get.

W2

W1

K2

K1

1g

a1

a2

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Dynamic Coupling

Now you can think the first system as a launcher and the

second system as a spacecraft. To minimize

response of the spacecraft, the spacecraft

should be designed to avoid dynamic

coupling with the launcher, i.e. designed

above the launch vehicle minimum

frequency requirement. Obviously the launcher and spacecraft are

more complicated than the two degrees

of freedom system. Coupled loads analysis

(CLA) is required to obtain the responses.

W2

W1

K2

K1

1g

a1

a2