What are key constraints for the spacecraft structure design?

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5/31/2007
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Pre-Class Assignment

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?

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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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• 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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• 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.

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
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 users 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 Planners 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
  users 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 users
  manual, e.g. gt20Hz 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 4.1 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
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 gt Supporting Plate gt Beam gt
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)
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
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 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
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, lets 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