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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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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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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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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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Reference
Spacecraft Systems Engineering, 2nd edition, Chapter 9,
Edited by Peter Fortescue and John Stark, Wiley
Publishers, 1995.
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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