衛星結構設計
祝飛鴻
10/26/2006
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Pre-Class Assignment:
1. What are the main functions of spacecraft structure?
2. What factors need to be considered for spacecraft
structure design?
3. What factors need to be considered on material selection
for space application?
4. What are the required major tasks for spacecraft structure
design?
5. How to verify spacecraft structure design?
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 What are the main functions of spacecraft structure?
 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
ARGO Satellite
- The first Taiwan designed satellite
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 What factors need to be considered for spacecraft structure
design?






Size
Weight
Field-of-view
Interference
Alignment
Loads
ARGO Satellite
- The first Taiwan designed satellite
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Falcon-1
Envelope
 Size:
 Fit into the fairing of candidate launch vehicle.
 Provide adequate space for component mounting.
30 cm
132 cm
13mm
clearance
123 cm
11mm
clearance
135 cm
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 Weight:
 Not to exceed lift-off weight of the selected launch vehicle.
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 Field-of-view (FOV):
 Define by other subsystems, e.g. attitude control
sensors, payload instruments, antenna subsystem, etc.
MSI FOV= 6 °
110 °
65 ° 65 °
110 °
X Band Antenna
FOV
Star Camera
FOV=  6.7° on short axis
 9.2° on long axis
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 Interference:
 With the launch vehicle fairing.
 Between components for physical contact
Falcon-1Envelope
and assembly.
GPS Ant.
8.6mm
clearance
Solar Panel
19mm
clearance
Section
Y=1219
X-Band Ant
15.5mm
clearance
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 Alignment:
 Define by other subsystems, e.g. attitude control sensors,
payload instrument, etc.
 On ground alignment.
 On-orbit thermal & hydroscopic distortion.
Requirement
Star
Camera
Orientation
± 0.5 (TBR)
Thruster
Orientation ±1.5 (TBR)
X-antenna
Orientation ±5
(TBR)
S-antenna
Orientation ±5
(TBR)
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 Loads:
 Environmental loads for structure design.
 Not-to-exceed loads for components and payloads.
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 What factors need to be considered on material selection
for space application?

Strength-to-weight ratio

Durability

Thermal stability

Thermal conductivity

Outgassing

Cost

Lead time

Manufacture
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 Commonly used material:
 Metals – Aluminum, Titanium, Magnesium, Beryllium
 Composites
 Ceramics
 Polymers
 Semiconductors
 Adhesives
 Lubricants
 Paints
 Coating
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Material Selection
Material
Density 
3
(Kg/m )
Young’s
Module E
(Gpa)
Yield
Strength S
(Mpa)
E/
S/ 
CTE
(m/m K)
Aluminum
7075 T6
2700
71
503
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186.3
23.4
Magnesium
AZ31B
1700
45
220
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129.4
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Titanium
Ti-6Al-4V
4400
110
825
25
187.5
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Beryllium
S 65 A
Fiber
Composite
- Kevlar
- Graphite
2000
304
207
152
103.5
11.5
1380
1640
76
220
1240
760
55
134
898.5
463.4
-4
-11.7
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 What are the required major tasks for spacecraft structure
design?
 Configuration design
 Environmental loads
 Structure analysis
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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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Configuration Design - ARGO
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Configuration Design
Structure
Configuration
Hardware
List
Mechanical
Layout
Hardware
Size
FOV
Requirements
Orientation
Requirements
Structure
Analysis
Environmental
Loads
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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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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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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 PPOD Payload Planner’s Guide” Revision C – June 5, 2000
for definition of launch loads.
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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
a
W
excitation with 3% damping ratio will
have 16.7g response at their natural
K
frequency, i.e.
For system 1: 16.7g at 40Hz
For system 2: 16.7g at 75Hz
1g
(Please refer to any vibration text book for derivation of results)
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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
a1
W2
K2
a2
W1
where 39.8Hz and 75.4Hz are the natural
frequencies of the combined system.
(Please refer to advanced vibration text book
K1
1g
for derivation of results)
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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
a1
W2
K2
a2
W1
where 38.3Hz and 41.8Hz are the natural
frequencies of the combined system.
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
a1
W2
system will increase from 23.2g to 99.2g.
K2
This phenomenon is called “dynamic
coupling”. The more closer natural
frequencies of the two systems, the
a2
W1
K1
higher response the system will get.
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
W2
a1
should be designed to avoid dynamic
coupling with the launcher, i.e. designed
K2
above the launch vehicle minimum
a2
W1
frequency requirement.
K1
 Obviously the launcher and spacecraft are
more complicated than the two degrees
1g
of freedom system. Coupled loads analysis
(CLA) is required to obtain the responses.
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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 Vehicle
Model
Combined
Model
Dynamic
Analysis
Spacecraft
Responses
Forcing Functions
of
Critical Flight Events
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Structure Analysis
 Once the mechanical layout is completed, the structural
analysis can be started. Major items include:
 Mass property analysis
 Structure member and load path
 Material selection
 Dynamic and Stress analysis
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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
W1
calculated based on the mass property
Y
of each individual elements e.g.
X
components, structure, hardness, etc. W2
D1
 The main purpose of mass property
analysis is to assure the design satisfies
D2
the weight and CG offset constraints
from the selected launcher.
Total Weight ?
MOI about Z axis ?
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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
Spacecraft Weight (lb)
1200
1400
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Structure Member 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.
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Structure Member and Load Path
Components => Supporting Plate => Beam => Supporting Points
Beam
Plate
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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.
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Dynamic & Stress Analysis
 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
-1  0
M. S. =
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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Dynamic & Stress Analysis
 Construction finite element model of a spacecraft is not
an easy 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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Dynamic & Stress Analysis
Structure
Configuration
Mechanical
Layout
Preliminary
Analysis/Design
Material
Selection
Quasi-Static
Loads
Load Path
Check
Approximation
Sizing
Preliminary
CLA
Detailed
Analysis/Design
Finite Element
Model
Final
CLA
Design
Verification
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 How to verify spacecraft structure design?
 Mechanical Layout – Assembly and integration
 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
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Homework Problem
 Describe future technology trend for spacecraft structure
design:
 Technology
 Goals
 Applications
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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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