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Space Mission Structures: From Concept to Launch

July 2013 Copyright Instar Engineering and Consulting, Inc.• instarengineering.com Materials may be reproduced in complete form only, with header and footer

Course developed and taught by Tom Sarafin and Poti Doukas

Course Sampler

Space Mission Structures From Concept to Launch

Objectives:

Audience: All mechanically inclined engineers and managers involved

in specifying, designing, building, and testing spacecraft or

launch vehicles and their components

Strengthen your understanding of …

– how structures fail

– how to design structures to be dependable for space missions

– how to reduce total cost and avoid problems through effective engineering

Make you think!

Overall objective:

Help you become

a better engineer!


Course Sampler


SMS History

First version developed by Tom Sarafin of Instar in 1995.

– Based on the book Spacecraft Structures and Mechanisms: From Concept to Launch (SSAM)

Book development jointly funded by Martin Marietta and many U.S. government organizations

4-year project, 24 authors, published in 1995 by Microcosm, Inc.

Tom Sarafin was the editor and principal author

– Original title of the course was the same as that of the book; taught 24 times under this title

Course revised in 1999; mechanisms dropped and new title, “Space Mission Structures: From Concept to Launch”.

– Taught 54 times under this title as of July 2013

The course is constantly being revised for improvement

– Based on in-class observations and new acquired knowledge

– Poti Doukas joined Instar in 2006 and has contributed to the improvement

The full version is offered as a 5-day (40-hour) course

– 32-hour version also available with selected topics omitted

Each participant receives a copy of the SSAM book


Course Sampler


Agenda

Introduction

1. Overview of Space Mission Structures

2. Review of Statics and Dynamics

3. Launch Environments and How Structures Respond

4. Mechanics of Materials

5. Strength Analysis

6. Structural Life Analysis: Fatigue and Fracture Mechanics

7. Overview of Finite Element Analysis

8. Preliminary Design

8a. Avoiding Problems with Loads and Vibration*

8b. Improving the Loads-cycle Process*

9. Designing for Producibility

10. Verification and Quality Assurance

11. A Case Study: FalconSAT-2

12. Final Verification and Risk Assessment

Summary/Wrap-up

*Sections 8a and 8b are included only in the 5-day, 40-hour version of this course.


Course Sampler


Everything Solid Is a Structure

Primary Structures

Tertiary Structures: • Brackets • Electronics Boxes

• Body Structure • Launch Vehicle Adapter

Secondary Structures

• Appendages • Solar Panels • Antenna Dishes • Support Trusses • Platforms

• PC boards

• What we normally refer to as a structure is something whose main

function is structural.

• But, regardless of function, everything made of solid materials—

radiators, feed lines, wires, microchips—is structurally loaded by

acceleration (e.g., during launch).

1-8


Course Sampler


A material can take only so much stress before failure occurs:

– Rupture

– Collapse

– Permanent deformation that is detrimental to function

How Launch Environments Affect Structures

Steady-state loads cause uniform acceleration, with a resisting quasi-static inertia load that stresses the materials.

Time-varying loads not only cause global acceleration but also cause structures to vibrate, which in turn stresses the materials.

Thrust

Turbulence

and wind

shear

Random

pressure

oscillation

from sound

Engine

vibration

Steady

winds Ultimate

failure

Yield

failure

The two structural characteristics

required of all hardware:

Strength = the highest load a structure can withstand (or highest stress a material can withstand) without failure

Life = cycles of load (or duration of load) before failure

1-10


Course Sampler


Weight Is Always a Consideration but Is Not Always Critical

Example: Estimated spacecraft weight (with growth allowance) = 1700 lb

Payload capacities at our target orbit for available launch vehicles:

LV Capacity

A

B

C

2000 lb

1500 lb

1200 lb

If we can’t drive our spacecraft weight down to

1500 lb (with an appropriate growth allowance),

then we must use launch vehicle A—and weight

will not be critical (300 lb margin).

When weight margin is available, we can reduce the cost of hardware

development and also decrease risk:

• Aluminum instead of advanced composites

• Proven designs that may not be as efficient

• Less costly manufacturing processes

• Robust designs (lower risk, simpler analyses and tests, etc.)

Myth: “All flight hardware is weight-critical.”

Reality: Often not true for launch-vehicle payloads

Make sure you understand how important weight is on your program!

Note: Mass properties may be critical for reasons other than launch, such as satellite on-orbit operation.

1-14


Course Sampler


Before we move on to dynamics,

Remember: Weight Is Not the Same As Mass!

maF

Force Mass Acceleration

mgw

Weight

(force) Acceleration of gravity

Weight

Acceleration

Mass

English units SI units

lb

in/s2

lb-s2/in

N

m/s2

N-s2/m (or kg)

Weight is in

units of force.

Kilogram is a

unit of mass,

not force.

Pound is a unit

of force, not

mass.

The weight of an object is less on the moon than on Earth, but

the mass is the same. It takes just as much force on the moon to

achieve a given acceleration for an object.

2-12


Course Sampler


Response to Sinusoidal Vibration: Transmissibility

Transmissibility describes the gain in a sinusoidally base-driven single-DOF system: peak response acceleration relative to peak input acceleration.

222

2

21

21

nn

n

f

f

f

f

f

f

TR

2

1QQuality factor,

= response ratio at resonance

f = Forcing frequency in Hz

fn = System’s natural frequency in Hz

= Damping ratio

(SSAM Eq. 5.30) where

Response to a sinusoidal force

depends on the damping and the

ratio of the forcing frequency to

the system’s natural frequency.

Ratio of forcing frequency to

natural frequency, f/fn

Transmissibility,

TR

(ratio of peak

response

acceleration to

peak base

acceleration)

SSAM Fig. 5.11

Amplification Attenuation (isolation)

Resonance

= 0.01

= 0.05

= 0.1

= 0.2

= 0.5

2

2-27


Course Sampler


Describing Random Vibration with Acceleration Spectral Density

Acceleration spectral density* (ASD) at frequency f, is the mean-square acceleration within a selected frequency band (whose center is f) divided by the bandwidth.

– Traditional units: g2/Hz

An ASD curve typically spans 20 to 2000 Hz for random vibration testing.

For easy control during test—and to account for uncertainty—we smooth out the ASD plot to envelop any significant peaks.

*Often referred to as power spectral density (PSD), but the

word “power” comes from processing an electrical signal

Frequency, Hz

Accel.

Spectral

density,

g2/Hz

10 100 1000 2000

1

0.1

0.01

0.001

Example:

Derived from

measured data

(hypothetical, in

this case)

Max expected

(acceptance test)

(typically 95%

probability, 50%

confidence)

Envelope of

measured data

Qualification test, 6 dB up per military standards

Protoqual test, 3 dB up per

military standards

The RMS (root mean square) acceleration is the square root of

the area under the ASD curve and is equal to the standard deviation

of random acceleration.

A 6-dB increase means the

acceleration doubles (g2

increases by a factor of 4).

Input Response

Any point on a vibrating component will have its own ASD,

different than the input ASD.

3-23


Course Sampler


Deriving A- and B-basis Allowables from Test Data

MMPDS Sec. 9 defines a rigorous process for deriving allowables as a standard for inclusion in MMPDS

When deriving allowables for internal use, one simple approach is to assume the strength data follow a normal distribution and to use the one-sided tolerance limit factors, k, based on sample size, n, from MMPDS Table 9.10.1:

n

.nln..exp.k

n

.nln..exp.k

193520095802821

87352203413262

90

99A basis:

B basis:

n k99 k90

6 5.18 3.03

8 4.42 2.60

10 4.02 2.36

12 3.77 2.22

15 3.53 2.07

20 3.30 1.93

25 3.16 1.84

30 3.06 1.78

100 2.68 1.53

300 2.52 1.42

2.33 1.28

Example:

10 specimens tested (n = 10)

Sample mean strength = 41.22 ksi

Unbiased sample standard deviation = 2.87 ksi

ksi 4348723622241

ksi 7298720242241

basisB

basisA

....F

....F

Allowable stresses:

4-7


Course Sampler


Understanding the Effect of Discontinuities

We make 100 identical tensile

specimens out of material XYZ

and pull them to failure.

Mean ult. strength = 34,297 lb,

Standard deviation = 1324 lb.

What would you expect

the mean strength to be?

2.00" 2.00"

0.25"

P

P

We then drill a 0.20”-dia hole

through each of 100

otherwise identical specimens

of material XYZ and pull them

to failure.

2.00" 2.00"

0.25"

P

P 0.20“-dia hole

4-9


Course Sampler


The Benefit of Ductility for Uniaxial Stress

Ductility is desirable because it allows materials to tolerate stress concentrations, which occur at discontinuities.

Example—Aluminum flat plate, with hole, under tension:

2.00"

P

P

0.20" dia.

Thickness, t = 0.25”

(0.635 cm)

s g = P

2.00 (0.25) = 2P

s max 3 s g

Net-section stress:

Stress, s

(ksi)

70 67

Strain, e 0.022 0.06

Gross stress:

Peak stress (elastic):

P (2.00 - 0.20)(0.25) n s = = 2.22P

= 6P When the edge of the hole is at a

rupture strain of 0.06, the strain at

the edge of the part is 0.022, and

the state of stress is nearly uniform.

Stress concentrations such as these often do not significantly reduce the

strength of parts made of ductile materials.

Such stress concentrations do, however, greatly reduce fatigue life.

Note: 1 ksi = 1000 psi = 6.895 MPa

SSAM Fig. 8.1

4-10


Course Sampler


What’s Wrong with This Design?

P

4-46


Course Sampler


Understanding Stress Analysis

Stress analysis: – Often thought of as the process of predicting stresses caused by

applied loads

– Actually is a process of relating applied loads to allowable loads (or allowable stresses, more often) in an apples-to-apples comparison

– Allowable stresses are derived from tests, but the stress itself is not measured

Load is measured and then converted (by some process) to stress

An apples-to-apples comparison means two things:

1. The structural design and failure mode of concern correspond

to those that were tested to derive the allowables

2. We use a method of converting load to stress that is consistent

with the way in which the allowable stress was derived

To the engineer, stress analysis is not about predicting stresses.

It’s about avoiding failure!

5-8


Course Sampler


A. Identify all potential

ultimate failure modes

of concern.

B. Explain how you would

assess those failure

modes and obtain

corresponding

allowables.

C. How would you improve

the design?

Class Exercise: Recognizing Potential Failure Modes (Problem 2)

Sandwich construction:

• Graphite/epoxy laminate face sheets

• Flexible aluminum honeycomb core

Uniformly

introduced

tension

and shear

at each

end

Resultant load

P

P

5-31


Course Sampler


Fatigue Damage is Cumulative

To assess fatigue life, most engineers use Miner’s Rule: ni = Number of loading cycles at stress level Si

Ni = Number of cycles to failure at stress level Si

j = Number of different stress levels

The number of cycles to failure at a given stress level comes from an S-N curve (stress vs. number of cycles), which is different for each material, stress ratio, and stress concentration factor.

An S-N curve is a “best fit” curve drawn through scattered test data. To account for variation in material life, multiply the predicted number of loading cycles, ni, by a scatter factor (life factor) of 4 (typical criterion).

Stress ratio,

Criterion:

Endurance limit = the stress at which a

material can withstand infinite cycles

(Aluminum alloys do not have a true

endurance limit.)

j

i i

i

N

nD

1

max

minRs

s

1 D

Kt = stress concentration factor

smin = minimum stress

smax = maximum stress

Cumulative damage,

SSAM Fig. 9.1

0

40

80

120

160

Steel 17-4 PH (H900) bar, longitudinal direction

(derived from MIL-HDBK-5E; not intended for reference) 200

Kt = 1.0 R = 0.1

Kt = 1.0 R = -1.0

Kt = 3.0 R = 0.1

Kt = 3.0 R = -1.0

Number of Cycles to Failure, N

10 3 10 4 10 5 10 6 10 7 10 8

where

Ma

xim

um

Str

ess,

sm

ax (

ksi)

6-4


Course Sampler


Example

• We’re designing a spacecraft structure with an octagonal cross section

What is wrong here?

Make sure the results pass the sanity check!

We have the FEM plot the mode shape, with colors to

indicate the associated stresses, and this is what we get:

Bolt circle

(grounded in model)

The color red

indicates the

highest stress

• The first mode of vibration is rocking on the separation mechanism

7-20


Course Sampler


FireSat Example

Mission: detect forest fires in the United States

Orbit: 700 km at 55° inclination*

Payload: mid-range infrared scanner

Payload field of regard: half angle of ±58° from nadir

Control system: 3-axis (because of off-nadir viewing)

Communications: relay satellite—Tracking and Data Relay Satellite System (TDRSS)

Launch vehicle: Pegasus

Allowable payload weight for 700-km altitude: 635 lb (288 kg mass)

Problem:

Develop a conceptual configuration for FireSat, a hypothetical satellite

conceived in Space Mission Analysis and Design (Larson and Wertz, ed.)

Preliminary Requirements:

*Inclination = angle between the satellite’s orbit plane and the planet’s equatorial plane

Example by John Leritz (SSAM Sec.14.4)

Pegasus data and constraints

used in this example are circa

1993 and may no longer apply.

8-4


Course Sampler


Designing Minimum-Weight Trusses

The work done by an applied force (external work) equals the

structure’s total strain energy (internal work).

For a truss made of n members with pinned ends, the total strain energy is

Pi = member i’s axial load

Li = length

Ai = cross-sectional area

Ei = Young’s modulus The work done by a single applied force, P, is

= displacement in the direction of applied force

Setting the external work equal to the internal work,

Thus, the way to stiffen a truss is to reduce its strain energy, which we can do by

increasing the members’ cross-sectional areas (which adds weight) or by making

load paths more direct (decreasing Pi2Li, which reduces weight).

Given optional truss arrangements with each member sized to provide just the

required strength for a given material, the lightest truss will be the stiffest.

n

i ii

ii

EA

LPU

1

2

2

2

PW

n

i ii

iin

i ii

ii

EA

LP

PEA

LPP

1

2

1

2

2

2or

22

SSAM Eq. 15.4

8-41


Course Sampler


Truss Example: Which Is the Most Efficient Arrangement?

Example by Bob Heymans

P

2P

2P

P

0

P

2P

2P

P

0

P

2P

2P

P

0

P

2P

2P

P

0 P

2P

-2.2P

0

P

2P

2P

P

0 2.2P

-2P

0

0

A

B (more efficient)

C

D

E

F

2P P

-1.4P -1.4P P

0 -P

P

0

0 P

1.4P 1.4P

-P -2P

-P

P

P

-1.4P 1.4P

-2P 0

P

2P

2P

P

0

2P 0

0 1.4P -1.4P

-P

20” (0.508 m)

(0.254 m) 10”

SSAM Fig. 15.14

(more efficient) (more efficient)

8-42


Course Sampler


Uncertainty in final weight

Weight margin is the difference between the allowable weight and the design weight. It’s good to have margin because your weight-growth allowance is probably based on average historical growth, and about 50% of past spacecraft have had more than average weight growth!

Understanding Weight Growth and Weight Margin

Weight-growth allowance (contingency) is the weight added to your current best estimate to account for uncertainty and omissions (has averaged about 25% overall for spacecraft between proposal and completion [ref. Hawkins*]).

Weight inevitably grows during the course of the program ...

SSAM Fig. 14.26

Margin

Estimated weight at conceptual

design

Launch vehicle capability (payload weight)

Weig

ht

Conceptual Design

Production and Test Launch

Program Phase

Expected weight growth

Design weight

(mature weight) =

estimated weight

multiplied by growth

factor

Growth factor =

1 + growth allowance

(decreases as the

program progresses)

* Results of survey of 16 commercial and military S/C. K. Hawkins, “Space Vehicle and Associated Subsystem Weight Growth”, Paper # 1816, presented at the 1988 Conference of the Society of Allied Weight Engineers.

8-54


Course Sampler


The Cost of Parts

The “after-the-fact” cost

study on the Saturn S-IVB

forward skirt was expanded

to compare expected costs

of alternate designs

Ref: O.P.Harwood, “Right for Flight:

The Structural and Architectural Design

of Machines that Fly”

As built

(A composite

skirt was also

considered in

the study but

had the same

weight as the

isogrid design

at a cost ratio of

9.3.)

9-6


Course Sampler


Example of Adaptability

Example of the rework necessary to

mount equipment in a skin-stringer

structure: S-IVB skirt. The rows of rivets

attach external stringers.

(Ref: O.P.Harwood, “Right for Flight: The Structural

and Architectural Design of Machines that Fly.”)

An integrally machined isogrid skirt provides

an attachment point at each node, allowing

the structure to adapt to late-added

components.

9-8


Course Sampler


Machinability

A metal’s machinability is a measure of the affordability of machining it, considering

– The time it takes to remove material

– The wear on the cutting tools

Inconel

Titanium

17-7PH

304 Stainless

303 Stainless

7075-T6

2024-T4

6061-T6

Relative Machining Time

The total cost of machining also includes preparation and setup time

SSAM Fig. 20.1

9-14


Course Sampler


Verification Can Be Proactive or Reactive

Reactive Methods

(weed out poor quality)

Analysis (when done after the

design is released)

Inspection

End-item testing:

•Qualification testing

•Acceptance testing

•Analysis-validation testing

Proactive Methods

(improve product quality)

Analysis (when done to improve a design)

Process control (goal is to learn to control

a manufacturing process so well that its

products do not need inspection)

Development test (to understand a

problem or to improve a design or

manufacturing process )

When production quantity is low, as it is for most aerospace programs,

we typically can’t afford to learn how to control processes as well as

we’d like, so we must rely somewhat on end-item inspections and tests.

The key is finding the right balance.

10-7


Course Sampler


Qualification and Acceptance Testing

Qualification test

Objective: Demonstrate the adequacy of a design by testing a dedicated non-flight article—built just like a flight article—to conditions that are more severe than expected for the mission. The additional factor or amount is the qualification margin, which is intended to cover build-to-build variation

Applications: When we are not confident in our ability to predict a design’s key characteristics or response or resistance to an environment

Limitations: Effectiveness depends on how well manufacturing processes are controlled (i.e., whether the qualification margin bounds the build-to-build variability)

Acceptance test

Objective: Demonstrate the adequacy of each product by testing it to conditions that are equal to or slightly more severe than expected for the mission

Applications: When we are not confident that our manufacturing processes will limit product variability within the range covered by the qualification margin

Limitations: Not effective in demonstrating life (no way of knowing how much life remains in the flight article)

A qualification random vibration or acoustics test provides confidence that

the flight units will have sufficient fatigue life after acceptance testing

10-17


Course Sampler


Hypothetical Problem (Example 12-1): Responding to a Negative Margin of Safety

• A spacecraft is about to complete integration and test, two months before launch.

• The verification loads cycle shows increased loads, causing a negative 20% margin of

safety (–0.20) for ultimate failure of a key member in the primary structure.

– If the member fails, the mission will be lost. (Human safety is not at issue.)

• All unnecessary conservatism has been scrubbed from the stress analysis and the loads

analysis; the analysis is relatively straightforward and the failure mode is well understood.

• The ultimate factor of safety used in the analysis is 1.25. Limit load is estimated at 3-sigma

probability (99.87%), and the allowable stress is A-basis (99%).

• The structure was tested, but to 1.1 times the original limit loads, which were 35% lower

than the new limit load.

• This is a one-of-a-kind structure; there is no qualification unit we can test.

• Redesign is out of the question—replacing the structure with a redesigned one would cause

the program to miss its launch window.

• Structural reinforcement could be added. Total estimated cost: $200,000 (labor, materials,

overhead)

• Mission value = $100,000,000 What would you do?

12-7


Course Sampler


Margin of Safety Relates to Probability of Failure

Load Strength

Negative margin of safety

• Doesn’t meet our criteria

• Implies greater risk

Load Strength

Zero margin of safety

• Acceptable

• What we want when weight is critical

F.S.

F.S. = factor of safety

Load Strength

Positive margin of safety

• More robust: better when weight is

not critical

F.S. M.S.

M.S. = margin of safety

12-15

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