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衛星結構設計

祝飛鴻5/25/2006

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What are the main functions of structure subsystem?

Provide support all other subsystems and attach the spacecraft to launch vehicle.

What factors need to be considered for structure design? Size Weight Field-of-view Interference Alignment Loads

ARGO Satellite- The first Taiwan designed satellite

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Size: Fit into the fairing of candidate launch vehicle. Provide adequate space for component mounting.

123 cm

135 cm

132 cm

30 cm

Falcon-1Envelope

13mm clearance

11mm clearance

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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.

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

and assembly. Falcon-1Envelope

SectionY=1219

Solar Panel19mm clearance

X-Band Ant15.5mm clearance

GPS Ant.8.6mm clearance

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

payload instrument, etc. 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 are major tasks to be performed for structure design?:

Configuration design

Environmental loads

Structure design and analysis

Design verification

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To accommodate all the components in a limited space while

satisfying its functional requirements, every spacecraft will

end up with a unique configuration.

Configuration Design

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Configuration Design - ARGO

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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”.

Environmental Loads

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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 P-

POD 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 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

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Structure Design & Analysis

Once the mechanical layout is completed, the structural

design and 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

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

Total Weight ?

MOI about Z axis ?

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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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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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PlateBeam

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

Structure Member and Load Path

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Material Selection

From purely structure design point of view, it is always

desirable to use material with high stiffness, high

strength, and low density, i.e. high strength/stiffness to

weight ratio. However, other factors may affect the

material selection, e.g. thermal conductivity, CTE

(coefficient of thermal expansion), cost, manufacture,

lead time, stability, etc.

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Material Selection

Material Density

(Kg/m )

Young’sModule E (Gpa)

YieldStrength S (Mpa)

E/ S/ CTE(m/m K)

Aluminum

7075 T6

2700 71 503 26 186.3 23.4

Magnesium

AZ31B

1700 45 220 26 129.4 26

Titanium

Ti-6Al-4V

4400 110 825 25 187.5 9

Beryllium

S 65 A

2000 304 207 152 103.5 11.5

Fiber Composite - Kevlar - Graphite

1380 1640

76 220

1240 760

55 134

898.5 463.4

-4 -11.7

3

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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 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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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.

Dynamic & Stress Analysis

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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Design Verification

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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Future Challenge

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Homework Problem

Derive a complete structure development process charts.

Input Input

Step 1

Output

Step 2

Output

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References

Spacecraft Systems Engineering, 2nd edition, Chapter 9,

Edited by Peter Fortescue and John Stark,

Wiley Publishers, 1995.