衛星結構設計
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衛星結構設計. 祝飛鴻. 5/31/2007. Pre-Class Assignment. What are key constraints for the spacecraft structure design? How the structure design is affected by other subsystems? How the structure design affects the performance of other subsystems?

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5830917

衛星結構設計

祝飛鴻

5/31/2007


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Pre-Class Assignment

What are key constraints for the spacecraft structure design?

How the structure design is affected by other subsystems?

How the structure design affects the performance of other subsystems?

How to distinguish a good and bad spacecraft structure design?


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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

Structure design

is affected by

all the other

subsystems

The first Taiwan designed satellite


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  • Spacecraft structure design has to satisfy the following factors:

    • Size

    • Weight

    • Field-of-view

    • Interference

    • Alignment

    • Loads

The first Taiwan designed satellite


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13mm

clearance

11mm

clearance

衛星尺寸限制:

  • Falcon 1 (Dia. 1371)

  • Falcon 1E (Dia. 1550)

  • Taurus-63 (Dia. 1405)

1.Size:

  • Fit into the fairing of candidate

    launch vehicle.

  • Provide adequate space for

    component mounting.

Taurus-63

Falcon 1E

Falcon 1

1371

1405

1550


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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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110 °

110 °

65 °

65 °

X Band Antenna FOV

3. Field-of-view (FOV):

  • Define by other subsystems, e.g. attitude control

    sensors, payload instruments, antenna subsystem, etc.

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.

GPS Ant.

8.6mm

clearance

Falcon-1Envelope

Solar Panel

19mm

clearance

X-Band Ant

15.5mm

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.


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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

Factors

to be

satisfied

Structure

Design

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

Combined

Model

Dynamic

Analysis

Spacecraft

Responses

Launch Vehicle

Model

Forcing Functions

of

Critical Flight Events


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Typical CLA Results


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Dynamic Coupling


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Structure Analysis

4. Structure Analysis:

4.1Mass 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

Y

W2

X

D1

D2


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Falcon-1 Launcher

Lateral CG centerline offset (in)

2.5

2.0

1.5

1.0

0.5

0.0

0 200 400 600 800 1000 1200 1400

Spacecraft Weight (lb)


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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.

Plate

Beam

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


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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 plate

with uniform loading)

Finite element solution

(Simply supported plate

with concentrated mass)

reaction

force

close form solution

(beam with concentrated 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 assessment

not 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 hasf2 = 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

a

K

1g


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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

a1

K2

W1

a2

K1

1g


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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

a1

K2

W1

a2

K1

1g


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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

a1

K2

W1

a2

K1

1g


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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

a1

K2

W1

a2

K1

1g


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