May 2009 

NASA/TM-2009-215747 

NESC-RP-05-126/04-093-E 

 

Deep Impact Delta II Launch Vehicle Cracked 

Thick Film Coating on Electronic Packages 

Technical Consultation Report 

Kenneth D. Cameron/NESC 

Langley Research Center, Hampton, Virginia 

Robert A. Kichak/NESC 

Goddard Space Flight Center, Greenbelt, Maryland 

Robert S. Piascik/NESC 

Langley Research Center, Hampton, Virginia 

Henning W. Leidecker 

Goddard Space Flight Center, Greenbelt, Maryland 

Timmy R. Wilson/NESC 

Langley Research Center, Hampton, Virginia 
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May 2009 

NASA/TM-2009-215747 

NESC-RP-05-126/04-093-E 

 

Deep Impact Delta II Launch Vehicle Cracked 

Thick Film Coating on Electronic Packages 

Technical Consultation Report 

Kenneth D. Cameron/NESC 

Langley Research Center, Hampton, Virginia 

Robert A. Kichak/NESC 

Goddard Space Flight Center, Greenbelt, Maryland 

Robert S. Piascik/NESC 

Langley Research Center, Hampton, Virginia 

Henning W. Leidecker 

Goddard Space Flight Center, Greenbelt, Maryland 

Timmy R. Wilson/NESC 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration. 
NASA Engineering and Safety Center 

Technical Consultation Report 

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Film Coating on Electronic Packages Technical 

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Deep Impact Delta II Launch Vehicle 

Cracked Thick Film Coating on Electronic Packages 

Technical Consultation Report 

Performed by 

The NASA Engineering and Safety Center (NESC) 

November 3, 2005
NASA Engineering and Safety Center 

Technical Consultation Report 

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NESC Request No. 04-093-E 

Table of Contents 

Volume I: Technical Consultation Report 

1.0 Authorization and Notification.........................................................................................4 

2.0 Signature Sheet...................................................................................................................5 

3.0 List of Team Members, Ex Officio Members, and Others.............................................6 

4.0 Executive Summary...........................................................................................................7 

5.0 Consultation Plan...............................................................................................................9 

6.0 Description of the Problem, Proposed Solutions, and Risk Assessment.....................10 

6.1 Mission Background......................................................................................... 10 

6.2 Technical Issue Description ............................................................................. 11 

6.2.1 Causal Factors .................................................................................................. 12 

6.3 NESC Risk Assessment.................................................................................... 13 

7.0 Data Analysis....................................................................................................................20 

7.1 Detailed Technical Discussion ......................................................................... 20 

7.1.1 Assessment of Material Properties................................................................... 21 

7.1.2 Environmental Crack Growth Properties......................................................... 23 

7.1.3 Assessment of Cracked TFA Thick Coating.................................................... 24 

8.0 Findings, Observations, and Recommendations...........................................................29 

8.1 Findings............................................................................................................ 29 

8.2 Recommendations............................................................................................ 30 

9.0 Lessons Learned...............................................................................................................32 

10.0 Definition of Terms..........................................................................................................32 

11.0 Minority Report...............................................................................................................33 

Volume II: Appendices

Appendix A. NESC ITA/I Request Form (NESC-PR-003-FM-01)............................................ 34 

Appendix B. References .............................................................................................................. 38 

Appendix C. Cutbar Essay........................................................................................................... 39 

Appendix D. List of Acronyms.................................................................................................... 40 

List of Figures

4.0-1 Deep Impact lifts off from pad 17-B at Cape Canaveral Air Force Station, Florida 

January 12, 2005 ................................................................................................................. 8 

6.1-1 Deep Impact Launch Configuration.................................................................................. 11
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6.2-1 TFA KSC S&MA Risk Chart ........................................................................................... 12 

6.3-1 KSC S&MA Deep Impact Residual Risk Assessment Summary..................................... 14 

6.3-2 LSP Risk Definitions........................................................................................................ 15 

6.3-3 LSP Risk Definition Matrix.............................................................................................. 16 

7.1-1 Photographs show TFA Thick Film Cracking.................................................................. 20 

7.1-2 Crack Velocity Curve for Sapphire in both Moist Air and Ringer’s Solution.................. 24 

7.1-3 Crack Velocity Curves for Sapphire Exposed to Air showing Increased Crack Growth 

Velocity with Increased RH.............................................................................................. 25 

7.1-4 Comparison of the Environmental Fatigue Crack Growth Characteristics of Sapphire 

in Moist Air and Ringer’s Solution (NaCl, Na lactate, KCl, CaCl).................................. 26

7.1-5 Pair of Cracks at ~50X of a Slab of Alumina (V-shaped Cracks).................................... 28 

7.1-6 Ratio of Figure 7.1-5: Closer View of Crack Image..........................................................28
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NESC Request No. 04-093-E 

VOLUME I: Technical Consultation Report 

1.0 Authorization and Notification 

Bryan O'Connor, Office of Safety and Mission Assurance (OS&MA), requested that the NESC 

provide a materials expert to evaluate two differing risk assessments regarding the start and 

propogation of cracks associated with thick film coatings on electronic packages (E-packages) on 

the Delta II Launch Vehicle. 

Mr. Ralph Roe, Director of the NESC, initiated the request for consultation on December 23, 

2004. 

The NESC approved the white paper at the NESC Review Board (NRB) on January 6, 2005. 
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2.0 Signature Sheet

NESC Technical Consultation Team

__________________________ _________________________________

Kenneth Cameron, NESC Lead Robert Kichak, NESC, GSFC

__________________________ __________________________________

Robert Piascik, NESC, LaRC Henning Leidecker, GSFC

___________________________

Tim Wilson, NESC, KSC
NASA Engineering and Safety Center 

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3.0 List of Team Members, Ex Officio Members, and Others 

Ken Cameron, NESC Lead Deputy Director for Safety 

Robert Piascik, NESC Materials Discipline Expert 

Robert Kichak, NESC Power & Avionics Discipline Expert 

Henning Leidecker, GSFC Chief Engineer, Parts, Packaging, and Assemblies Technologies 

Office 

Tim Wilson, KSC NESC Chief Engineer 

Consultants 

Derek Harris, KSC 

Raoul Caimi, KSC 

John Martin, KSC 

John Tinsley, KSC 

Bert Garrido, KSC 

Omar Baez, KSC 

Robert Abernathy, The Aerospace Corporation

Darren Bedell, Launch Services Program 

Steve Francois, Launch Services Program 
NASA Engineering and Safety Center 

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4.0 Executive Summary 

The Deep Impact spacecraft was launched on a Boeing Delta II rocket from Cape Canaveral Air 

Force Station (CCAFS) on January 12, 2005. Prior to the launch, the Director of the OS&MA, 

Mr. Bryan O’Connor, requested the NESC to lead a team to render an independent opinion on 

the rationale for flight and the risk code assignments for the hazard of cracked Thick Film 

Assemblies (TFAs) in the E-packages of the Delta II launch vehicle for the Deep Impact 

Mission. The concern was with the possibility of cracks and crack growth in the TFA in the 

servo-control electrical packages on the Delta II launch vehicle. This issue surfaced during the 

System Mission Assurance Readiness Review (SMARR). The KSC Launch Services Program 

(LSP) and the Boeing Launch Services Contractor (LSC) presented this item as no risk, and 

Kennedy Space Center (KSC) Safety and Mission Assurance (S&MA) presented it as a medium 

residual risk item. 

The NESC team concluded that the KSC S&MA, Boeing and the LSP all agreed on the nature 

and cause of the TFA cracking hazard, and on the existing rationale for flight. However, they 

used different rationales in assigning risk codes when using a common assessment tool. The 

NESC team concluded that the TFA crack risk should be considered to be in the range of 

medium-to-low risk, and represented an acceptable risk for flight. The NESC team concurred 

with KSC S&MA that “5x2” more accurately characterized this risk. Boeing and the LSP 

acknowledged the probability of failure to be “less than 1 percent”, but not zero, as assessed by 

their subjective methods. KSC’s S&MA and the NESC team agreed that the TFA cracking 

hazard represented a reportable risk for the Flight Readiness Review (FRR). The possible cause, 

excessive conformal coating under the TFA assemblies due to manufacturing changes, was 

identified, but was not verified by testing. An enhanced component-level screening test, which 

reviewed parametric data from box-level testing while looking for changes, has proven effective 

to-date. However, there is some risk that this technique may not be 100 percent effective 

because it is an indirect method of detection. Due to the potential for undetected micro-cracks to 

propagate with time or with the launch environment, a residual risk remained to trigger failure of 

an assembly with micro-cracks that are under stress. 

The NESC team concluded that Boeing and the LSP should continue to pursue other design 

approaches to further reduce the risk to TFA circuits in the future. 
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Figure 4.0-1. Deep Impact lifts off from pad 17-B at Cape Canaveral Air Force Station, 

Florida January 12, 2005 

(NASA photo) 
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5.0 Consultation Plan 

The NESC approved the white paper at the NRB on January 6, 2005. The mission of this effort 

was to investigate the possibility of cracks and crack growth in the TFA in the servo-control 

electrical packages on the Delta vehicle. After evaluation of the risks by the KSC assessment 

team and Boeing, it was realized that the means for rating risks varied between the organizations, 

making it difficult to confirm a rating assignment for the FRR. 
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6.0 Description of the Problem, Proposed Solutions, and Risk Assessment 

6.1 Mission Background 

The Deep Impact mission is a partnership with the University of Maryland (UMD), the 

California Institute of Technology's (CIT) Jet Propulsion Laboratory (JPL), and Ball Aerospace 

and Technology Corporation (BATC). The scientific leadership of the mission is based at UMD. 

Engineers at BATC designed and built the spacecraft under JPL's management. Engineers at 

JPL control the spacecraft after launch and send data to scientists for analysis. The mission is 

implemented with a two-part spacecraft. The larger "flyby" spacecraft carries a smaller 

"impactor" spacecraft to the comet Tempel 1, and releases it into the comet's path for a planned 

collision. 

In January 2005, a Delta II rocket (configuration shown in Figure 6.1-1) launched the combined 

Deep Impact spacecraft which, after leaving Earth's orbit, was directed toward the comet. The 

combined spacecraft approached the comet and collected images before the impact. In early July 

2005, 24 hours before impact, the flyby spacecraft pointed high-precision tracking telescopes at 

the comet and released the impactor on a course to hit the comet's sunlit side. 

After release of the impactor, the flyby spacecraft maneuvered to a new path that, at closest 

approach, passed 500 kilometers (km) (300 miles) from the comet. The flyby spacecraft 

observed and recorded the impact data, the ejected material blasted from the crater, and the 

structure and composition of the crater's interior. After its shields protected it from the comet's 

dust tail passing overhead, the flyby spacecraft turned to re-look at the comet. The flyby 

spacecraft took additional data from the other side of the nucleus and observed changes in the 

comet's activity. While the flyby spacecraft and impactor performed, professional and amateur 

astronomers at large, and small telescopes on Earth, observed the impact and its aftermath. 

Results were broadcast over the Internet. 
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NESC Request No. 04-093-E 

 

Vehicle Configuration

NCS System

Third-Stage Motor

Separation Clamp Band Spin Table

Guidance Electronics

Second-Stage Miniskirt

and Support Truss

Helium Spheres

Nitrogen Sphere

Interstage

Wiring Tunnel

Fuel Tank

Centerbody Section

Oxidizer Tank

Conical Section

Fairing Access Door

First Stage

Fairing

Thrust

Augmentation

Solids

Second Stage

Third Stage

Motor

Fairing

• Vehicle configuration: 7925-9.5

• Launch site: SLC-17B at CCAFS

• Launch period: 30 Dec 2004 - 28 Jan 2005

• Unique mission requirements:

• 3712C payload attach fitting

• Third stage motor dome blanket

(with metal Velcro)

• Nutation control system

(37 lbf reaction engine assembly)

• De-spin system

• Third-stage ballast (10 ± 10 lb)

• 2 mission-specific access doors

• Two 37-pin PAF umbilicals

Deep Impact

Spacecraft

Attach Fitting

Figure 6.1-1. Deep Impact Launch Configuration 

6.2 Technical Issue Description 

At the S&MA FRR, the issue of electronic package (E-package) cracked TFAs was assessed by 

the KSC S&MA organization as a “5x2” residual risk. Note that the Boeing and the LSP did not 

list this item as a reportable risk. The issue involved a number of E-packages that contained 

TFAs that were manufactured at an El Paso, Texas facility following a supplier change from 

Monrovia, California (this supplier change resulted in process changes). Cracks were found in 

some TFAs in El Paso during 2002. An enhanced screening technique was put in place to 

perform an enhanced review of box-level acceptance test data to look for parametric changes of 

greater than 5 percent as an indicator of possible TFA cracks. Appendix C shows the formula 

and the plot of DeltaR/R versus Crack Size1

. This issue also was previously addressed for the 

Swift and Messenger missions’ launch vehicles. Two of these potentially “suspect” E-packages, 

containing a total of twelve TFAs, were used in the Deep Impact Launch Vehicle: one E-package 

1

 Extraction from Henning Leidecker’s Cutbar Essay. 
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NESC Request No. 04-093-E 

with one circuit board with two TFAs in the first stage, and one E-package with two circuit cards 

with a total of eight TFAs in the second stage. These E-packages are part of the vehicle’s 

guidance navigation & control (GNC) system. TFA failure during launch can be catastrophic. 

The chart from the S&MA FRR is shown in Figure 6.2.-1, which summarizes their assessment of 

the issue. 

E-Package Cracked Thick Film Assemblies

RISK TYPE: Mission Success

HAZARD REPORT: N/A

CRITICALITY: N/A

RISK CATEGORY: Technical

ORGANIZATION: Safety and Mission Assurance

ASSIGNED TO: Raoul Caimi, KSC

INDEPENDENT ASSESSORS: SA-G

RISK DESCRIPTION:

A cracked TFA has been detected in a first stage E-pkg at the HY1 location. This is the first such departure from the HY3 location for 

this issue, as seen previously. The existing enhanced screening technique’s ability to detect the presence of such a crack was 

evaluated for applicability to the first stage E-pkg. This screen has been previously used for 2nd stage E-pkgs on MESSENGER and 

Swift. Both primary and secondary circuits can be affected by a TFA crack leading to loss of mission.

CAUSE:

Not determined. Boeing is attributing the cracks to conformal coating shrinkage during cure.

RISK EFFECTS:

Loss of mission.

RISK REDUCTION ACTIONS:

An enhanced screening technique has been developed to detect TFA cracks for second stage E-pkg’s. Boeing and LSP used the 

same technique to detect TFA cracks in first stage E-pkg’s. The Deep Impact 1st stage E-pkg passed the screening with no indication 

of a cracked TFA.

CONSTRAINTS TO FLIGHT:

None.

RECOMMENDATION / RATIONALE:

The screening technique detects the presence of existing detrimental cracks; therefore, this presents a risk that can be accepted for 

flight.

1 234 5

5

4

3

2

1

X

MISSION SUCCESS RISKS

IMPACT

P

R

O

B

A

B

I

L

I

T

Y

X: SMA Risk Assessment

1 234 5

5

4

3

2

1

X

MISSION SUCCESS RISKS

IMPACT

P

R

O

B

A

B

I

L

I

T

Y

X: SMA Risk Assessment

Figure 6.2-1. TFA KSC S&MA Risk Chart 

6.2.1 Causal Factors 

The TFAs contain circuit traces (conductors) on an alumina substrate, as well as surfacemounted electronics parts. Some of these parts include high precision resistors that are bias and 

gain setting elements for op-amps in servo amplifier circuits. Boeing hypothesized that the new 

process changes introduced the possibility of excess conformal coating material between the 

TFA modules and the polyamide printed circuit card to which they are mounted and connected. 

As the conformal coating material cures and shrinks, Boeing suggested that stresses can be 

induced in the alumina TFA module and can result in cracks. 
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Following the initial discovery of the problem in 2002, an “enhanced screening” technique was 

instituted where the box-level acceptance test data was reviewed for parameter changes, or 

deltas, between the initial Acceptance Test Procedure (ATP) and subsequent ATPs. A threshold 

for investigation was set to be a parametric change of 5 percent or greater from one ATP to 

another. The enhanced screening was a manual review of the detailed parametric data collected 

during the ATP. Cracks in the TFAs were detected by failures which propagated into parts on 

the boards, most notably high stability precision laser-trimmed resistors that set amplifier gain or 

biasing. These cracks manifested themselves in observed parametric shifts in the data during the 

detailed box ATP testing. Absence of parameter shifts between initial and final acceptance 

testing is considered to be evidence that an E-package that does not contain cracks of concern. 

These box-level ATPs were run between 201 and 500 days after the cards were originally 

conformal-coated. It was possible for the boxes to pass original go/no-go functional testing, but 

still have parametric shifts. 

The basis for flight or flight rationale is the understanding of the likely root cause of the failures, 

and passing the box-level enhanced parametric data screening. The residual risks are that there is 

less than 100 percent certainty that the true root cause has been identified and there are 

reasonable questions as to the effectiveness of the screening. 

6.3 NESC Risk Assessment 

The risk assessment process included the identification of hazards, assessment of hazard 

consequences, assessment of likelihood of hazard occurrence, and flight rationale and acceptance 

of the assessed risk. The differences of opinion between the KSC S&MA organization and the 

Boeing and the LSP were due to subjective assessments based on engineering experience, 

differences in optional approaches for the application of assessment tools, and translation 

between limited reliability data and system failure probability. 

The overall risk assessment performed by KSC S&MA is shown in Figure 6.3-1. The TFA crack 

issue was assessed as a medium residual risk item. 
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1 23 4 5

4

5

4

3

2

1 2, 3

1

MISSION SUCCESS RISKS

CONSEQUENCES

L

I

K

E

L

I

H

O

O

D

1 23 4 5

5

4

3

2

1

SAFETY RISKS

CONSEQUENCES

L

I

K

E

L

I

H

O

O

D

SMA 3- Microsemi Transistor M

SMA 4- RIFCA Alignment M

SMA 1- Thick Film Assembly M

SMA 2- GEM Delaminations M

Assurance Elements Program KSC SMA Minority Opinion

S = Safety

M = Mission Success

SMA 3- Microsemi Transistor M

SMA 4- RIFCA Alignment M

SMA 1- Thick Film Assembly M

SMA 2- GEM Delaminations M

Assurance Elements Program KSC SMA Minority Opinion

S = Safety

M = Mission Success

Figure 6.3-1. KSC S&MA Deep Impact Residual Risk Assessment Summary 

The LSP and the launch vehicle manufacturer used a risk assessment matrix as shown in Figures 

6.3-2 and 6.3-3. An assessment of probability of an issue causing failure within 100 flights was 

made based on the judgment of program engineering. 
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Risk Definitions - Probability

Probability of 

occurrence is 

very high. 

Existing 

processes and 

mitigation 

efforts cannot 

prevent this 

risk scenario; 

no alternative 

processes or 

mitigation 

efforts are 

available.

Probability of 

occurrence is 

high. Existing 

processes and 

mitigation 

efforts cannot 

prevent this risk 

scenario; a 

different 

process or 

mitigation effort 

might.

Probability of 

occurrence is 

moderate. 

Existing 

processes and 

mitigation 

efforts may 

prevent this risk 

scenario, but 

additional 

actions will be 

required.

Probability of 

occurrence is 

low. Existing 

processes and 

mitigation 

efforts are 

usually 

sufficient to 

prevent this risk 

scenario; 

additional 

actions may be 

required.

Probability of 

occurrence is very 

low. Existing 

processes and 

mitigation efforts 

are strong and 

very likely to 

prevent this risk 

scenario

Guidance for *1-5% 6-10% 11-50% 51-90% 91-<100%

Probability 

Level 

Selection

Label Very Low Low Moderate High Very High

Level 1 2 3 4 5

Probability Classification

Probability of 

occurrence is 

very high. 

Existing 

processes and 

mitigation 

efforts cannot 

prevent this 

risk scenario; 

no alternative 

processes or 

mitigation 

efforts are 

available.

Probability of 

occurrence is 

high. Existing 

processes and 

mitigation 

efforts cannot 

prevent this risk 

scenario; a 

different 

process or 

mitigation effort 

might.

Probability of 

occurrence is 

moderate. 

Existing 

processes and 

mitigation 

efforts may 

prevent this risk 

scenario, but 

additional 

actions will be 

required.

Probability of 

occurrence is 

low. Existing 

processes and 

mitigation 

efforts are 

usually 

sufficient to 

prevent this risk 

scenario; 

additional 

actions may be 

required.

Probability of 

occurrence is very 

low. Existing 

processes and 

mitigation efforts 

are strong and 

very likely to 

prevent this risk 

scenario

Guidance for *1-5% 6-10% 11-50% 51-90% 91-<100%

Probability 

Level 

Selection

Label Very Low Low Moderate High Very High

Level 1 2 3 4 5

Probability Classification

*Percentages are for comparative purposes only.

SMA does not use percentages shown

Source: Launch Services Program Risk Management Plan (KSC-PLN-2130)

Figure 6.3-2. LSP Risk Definitions 
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Risk Definitions

54321

1x1 2x1 3x1 4x1 5x1

1

1-5%

1x2 2x2 3x2 4x2 5x2

2

6-10%

1x3 2x3 3x3 4x3 5x3

3

11-50%

1x4 2x4 3x4 4x4 5x4

4

51-90%

1x5 2x5 3x5 4x5 5x5

5

91-<100%

54321

1x1 2x1 3x1 4x1 5x1

1

1-5%

1x2 2x2 3x2 4x2 5x2

2

6-10%

1x3 2x3 3x3 4x3 5x3

3

11-50%

1x4 2x4 3x4 4x4 5x4

4

51-90%

1x5 2x5 3x5 4x5 5x5

5

91-<100%

Impact

Probability 

of 

Occurrence

Source: Launch Services Program Risk Management Plan (KSC-PLN-2130)

SMA does not use the percentages shown

Figure 6.3-3. LSP Risk Definition Matrix 

Risk Discussion 

Until recent times, the term "risk" was equated with "hazard, danger, peril, exposure to loss, 

injury, or destruction". There was no sense that likelihood of this event actually happening was 

part of the term "risk". "Risk" only meant a possibility of incurring loss or misfortune, and did 

not address whether this possibility was likely or unlikely. "Risk" was a one-dimensional term. 

Only in the last few decades has "risk" become two-dimensional. The subject of "risk 

management" supposes that we can both quantify the loss associated with a "risk event", and also 

quantify the likelihood of this "risk event". A failure effects analysis is designed to help with the 

first quantification, and some sort of probability analysis is designed to help with the second. 
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A large practical problem, however, is that resources were never spent to collect the data needed 

to make a reliable probability analysis for some "risk events", especially for rare events that 

infrequently happen and happen in expensive systems. Therefore, in many cases, these 

probability evaluations were made as a guess. In particular, this is the case for the various 

probability evaluations for the cracking of sintered alumina TFA, used in the GNC system of the 

Delta II chosen for the Deep Impact launch. 

Boeing has offered an explanation for these cracks: the choice of conformal coating material and 

the manner of its application has caused stresses that have risen to the level of the fracture 

strength of the sintered alumina. Thus, some have cracked. This explanation is plausible; 

however, it has not been established with rigor. For example, it does not address the role played 

by cracks in the surface of these alumina plates (induced by handling and by the laser cuts used 

to trim the resistors on the surfaces of these plates). These surface cuts are not likely to be "all 

the same" as one moves from specimen to specimen. No fractrography seems to have been 

conducted to identify the source of each crack, especially for the ones that have completely 

shattered the alumina plates. The fraction of the cracked specimens has not been reproduced in 

any way by Boeing's explanation. 

It may be possible to regard these alumina plates as having been subject to a "proof test" by the 

pre-flight tests. This provides confidence that those alumina plates, that have survived so far, 

will not fail during the much less stressful launching conditions. However, Boeing has not 

presented this "proof test" argument in an orderly manner, using the standards now established in 

areas of fracture mechanics, which report quantitative studies of the effectiveness of different 

"proof test" designs for particular systems. 

Thus, the probability of success "P" is not rigorously known. The value of P or its uncertainty 

limits, "P minimum, P maximum", are not known in an objective manner. A reliable value for 

this P is not known, although it can be supposed it is more likely to be "low" (less than 10 

percent) than "high" (greater than 10 percent), and even more likely to be "very low" (less than 1 

percent) than simply low. 

The risk matrix used by US programs follows our legal system does not allow "not known" as a 

measure of the probability of failure. It compels one to choose a category. In response to this 

constraint, it is the policy of some organizations to assign "high risk" to the cases when one 

cannot objectively assign a probability value. This has happened in this circumstance where the 

KSC S&MA team has reacted to the uncertainty in the value of P with the finding that "this 

likelihood is appreciable": level two. Level two is also assigned to be consistent with the KSC 

S&MA’s experience base for large number of evaluations. 
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Boeing has reported that their engineers were surveyed with the question: "Based on your 

experience, and supposing one hundred launches, what is the number of launches that would fail 

as a result of using control systems like the ones considered for this Deep Impact launch?" The 

response was that this was judged to be less than one launch failure per one hundred launches. 

So, Boeing has concluded that the estimate of the probability of launch failure caused by a crackinduced fault in a TFA is less than 1 percent. This falls below the threshold at which Boeing 

reports risks. Boeing judged this risk to be "below the chart's lower edge". However, they did 

not mention whether their engineers agreed to the "less than 1 percent"; whether there were those 

who supposed a larger value, and only the average of all these estimates was below 1 percent; or 

whether there were some who declined to make any estimate. These all point to a possible lack 

of consistency with these engineering estimates. 

This discussion explains how the Boeing and the KSC S&MA teams have provided different 

reports as to launch risk: their policies are different and non-comparable. The elasticity of the 

term "probability", presently used on risk charts, extends to include these extremely different 

uses: engineering guesses on the part of Boeing versus an attempt to communicate uncertainty on 

the part of the KSC's S&MA team. Boeing's choice of "don't report if P < 1 percent" has an 

implication that is troubling. Most likely there are many hundreds of subsystems like the TFAs 

of interest. Assume this number as N. Suppose that each has a problem that presents a 

probability for launch failure of "f" that is just under 1 percent, so the probability of success is p 

= 1 - f = 99- percent. For the launch to succeed, each of these N subsystems has to work; the 

probability of a successful launch is then P = p**N. The table below illustrates values for launch 

success "P" for various levels of N: 

N P 

10 90 percent 

30 74 percent 

100 37 percent 

300 4.9 percent 

Thus, Boeing could be non-reporting a collection of problems, each falling below their "1 

percent" criteria, which could still dramatically lower the probability of a successful launch. The 

only way for a complicated system to be reliable is for each of its subsystems to be remarkably 

reliable. As soon as subsystems are accepted that are less than remarkably reliable, launch 

failures will occur. 

Thus, this investigation surfaces some differences in policy between assigning values to the 

"likelihood of failure", as used by the Boeing and S&MA teams. It also surfaces a reporting 

policy ("<1 percent" = do not report) on the part of Boeing that can have an ominous 
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consequence for systems as complex as a launch vehicle. Is there a way to arrive at a better 

estimate for the probability of failure caused by a crack in a TFA affecting its performance 

during launch, given the pre-testing that these units have had? Perhaps further exploring the 

"proof testing" approach may provide a more objective quantitative basis. 
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7.0 Data Analysis 

7.1 Detailed Technical Discussion 

To assess the performance of the Delta second stage TFA, it is important to understand the 

fracture properties of the sintered alumina (Al2O3) TFA coating. The alumina coating exhibits 

brittle fracture characteristics typical of most ceramic materials. This is evidenced by the classic 

brittle fracture shown in Figure 7.1-1. The following sections briefly describe the typical 

fracture properties expected for the TFA coating exposed to tensile load. In general, tensile stress 

rather than compressive stress promote cracking in these materials. Single crystal Al2O3

(sapphire) data is used to approximate the general fracture properties of the TFA. (The single 

crystal data is used here to establish likely trends and not absolute properties). 

Figure 7.1-1. Photographs show TFA Cracking 
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Note: Cracks exhibit brittle “glass-like” fractures. (Reference: NASA Presentation: “Cracked 

Thick Film Assemblies, Delta II, Stage 2 Servo Amplifier CCA EBR-02043, contributions 

by Thomas Bulk et al) 

7.1.1 Assessment of Material Properties 

For the TFAs, the stress has been applied by bending the specimen. This means that the largest 

stresses are, at the surface, on the outside of the imposed bend (i.e., the side that is stretched as 

the film bends), and so surface cracks on this surface are of dominating importance. (For 

practical purposes, the internal defects associated with voids left from incomplete densification 

during sintering can be ignored). These dominating surface cracks are not defined by knowing 

that this material is "sintered alumina", but rather on the handling that the specimen has been 

given. 

For example, sintered alumina plates do not have intrinsic fracture strength by the time they have 

reached service. Rather, the largest bending stress that they can sustain is entirely controlled by 

the presence of surface cracks installed by the handling they receive up until failing. This 

handling causes a distribution of cracks of various sizes "a". The largest of these, in a high-stress 

region, will dominate the observed fracture strength. 

It has long been known that glass tubes that have been handled (even by letting one tube touch 

another) have their resistance to breaking when bent sharply reduced. This strength can be 

restored essentially back to its "as new" value by polishing the surface (either by the use of 

rouge, by "fire polishing," or by etching in hydrofluoric acid). This illustrates that brittle 

materials do not have an intrinsic strength in the field, after handling, but have a strength against 

bending stresses that is entirely determined by the surface cracks installed by handling. 

One common way for a sintered alumina plate to acquire a highly damaging nick is for another 

plate to brush against it. Contact with hardened steel (such as a file or a scribe) is another way to 

scratch the surface of the alumina. Placing the sintered alumina plate onto some hard surface can 

induce edge-cracks. 

If the edges of these sintered plates are not chamfered, but are left in the form of right-angles, 

then edge-chipping is inevitable as the plates are handled. Typically, it may be taken for granted 

that any bit of optical glass such as a mirror, a window, or a lens will be chamfered. However, in 

E-Packaging applications, substrates of sintered alumina are typically not chamfered. Usually, 

these are hard-bonded to a strong backing and not subject to bending. But these TFAs are 

subject to bending stresses, and so they should be edge-chamfered. One of the pictures of the 

shattered plates seems to show that the cracks started from edges. The laser-induced cuts used to 

trim the TFA resistors are classic "cracks" into the surface of the alumina. Some researchers 
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studying the fracture properties of alumina have taken to using laser-cuts as an especially 

controlled way of inducing a well-defined crack into the surface of the alumina. 

It seems unlikely that each of the sintered alumina specimens comprising the substrate of these 

TFAs would have the same handling-induced damage. Therefore, it would seem that the family 

of them would show great variability in the time to fracture (one hundred to one variations are 

easy to understand), even when subject to the same applied stress. 

It is possible that the population of edge-cracks, and perhaps cracks on the surface away from the 

edges, play a role in the fracturing of the TFAs. This is important in understanding which plates 

fractured and which ones did not. A positive note is that the rate of crack growth, under applied 

stress, increases as rapidly as the applied stress is increased that "proof testing" can work if the 

specimen has passed a given load without fracture. Then, it is highly unlikely it will 

subsequently fracture under a substantially smaller load. Additional study of "C" and "p" for 

sintered alumina and these relationships (as described in the “Fracture Properties” section below) 

may be able to show a quantitative model of the degree to which the Boeing tests have acted as a 

"proof test". This would give quantitative support to Boeing's position that all the plates that are 

ever going to crack under launch loads have already been forced to crack by pre-launch stresses. 

This approach may be useful for future flights. 

Alumina Processing 

Al2O3 can be prepared as equi-axial crystals of a moderately uniform diameter and can be 

sintered into a solid by exerting pressure at a high temperature. The result is a material with a 

density approaching that of the single crystal form (and thus one parameter characterizing this 

sintered form is its density, while another is the grain diameter). A typical value is 98 percent 

used as an E-Packaging substrate (sometimes the value can be up to 99.8 percent). 

Fracture Properties 

The fracture mechanics of the sintered form is the subject many publications. Used as a 

packaging substrate, as it is in the TFAs of present interest, sintered alumina shows "brittle 

fracture". When subjected to an increasing stress, its strain increases linearly up to a certain 

threshold of stress and then the specimen shatters. There is no region of ductile deformation in 

which the sample plastically deforms. Hence, the critical stress that separates "safe" linear 

deformation from "catastrophic" rupture must be understood. 

Many studies show that the sintered alumina is suffused with cracks. One class of these is the 

micro-cracks surrounding internal voids left from incomplete densification during sintering. 

Another class is surface cracks of all sizes ― from atomic dimensions (always) to as large as 
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visible (sometimes) ― resulting from the formation and the handling of the specimen. Purely 

compressive stresses force all cracks to close. Alumina is remarkably strong under purely 

compressive stresses. 

Tensile stresses force all cracks, internal as well as surface cracks, to grow. The rate of growth is 

increased by the presence of catalytic agents such as water (present in the environment as 

moisture in the air, which is concentrated by surface tension at the tip of a crack into essentially 

liquid densities). One useful relation is crack growth rate = da/dt = C Kp

 where "a" is the length 

of the crack, "t" is the time, "K" is the stress factor, "C" is a pre-factor, and "p

" is an exponent. 

For some alumina specimens, values of “p” as large as 60 have been measured. The value of "C" 

for silica glass is roughly 105

 times larger in the presence of moist air than in an excellent 

vacuum. However, water has a much smaller effect on "C" for sintered alumina. The relation 

between the stress factor "K" and the applied stress is controlled by the crack length and shape 

and its orientation to the applied stress. This relation is defined in texts on fracture mechanics. 

Crack Growth Properties 

The TFA alumina will likely exhibit typical ceramic tensile properties such as high modulus, 

low total ductility, and low fracture toughness. These result in brittle crack growth and fracture. 

Typically, resistance to sustained load cracking is poor. Initially, stable crack growth rapidly 

becomes unstable and leads to fracture at relatively low tensile loads (low crack-tip stress 

intensity threshold - KTH=1.5 MPa√m) for single crystal alumina (sapphire) in humid air (shown 

in Figure 7.1-2). As shown, the near vertical slope of the "crack growth rate versus crack-tip 

stress intensity" curve is typical of the spectacular crack growth characteristics of these 

materials, which make them potentially dangerous in high-stress applications that can be ruined 

by cracking. 

A small change in crack-tip driving force (tensile stress) results in a dramatic increase in crack 

growth rate until unstable crack growth and final fracture occurs and fracture toughness is 

rapidly reached (K1C≈2.1 MPa√m). Similar brittle crack growth characteristics are observed 

under cyclic loading, also shown in Figure 7.1-2. These data illustrate that the TFA coating 

material is “crack sensitive”, which means that small flaws (micro-cracks) can readily propagate 

under tensile loading. 

7.1.2 Environmental Crack Growth Properties 

Alumina is susceptible to environmental-assisted cracking under sustained load, stress corrosion 

cracking (SCC) and cyclic loading, and corrosion fatigue (CF). Figures 7.1-2 and 7.1-3 show the 

SCC characteristics of sapphire exposed to different levels of relative humidity air and Ringers 

solution (NaCl, Na lactate, KCl, CaCl). Figure 7.1-2 also shows that what are commonly 
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thought to be relatively benign environments (50 percent relative humidity (RH) air) can increase 

crack growth rates in sapphire. These data show that time-dependent environmental crack 

growth damage modes can occur in Al2O3 materials and should be considered when cracking is 

possible. Figure 7.1-4 shows that under fatigue loading, a dramatic change in environmental 

crack growth characteristics is observed. Also depicted, a factor of 102

 increased fatigue crack 

growth rates is observed for the Ringers solution compared to humid air. Again, these data show 

that the environment can have deleterious effects on the crack growth rates of these materials. 

7.1.3 Assessment of Cracked TFA Thick Coating 

Given the brittle fracture properties of the TFA, it is likely that the rigorous ATP testing screened 

most damage. The “crack sensitivity” and rapid crack growth characteristic of these materials 

would likely produce failure during the ATP testing. Whether undetected damage was produced 

remains a valid point of discussion. 

KTH KIC

K (MPa√m)

KTH KIC

K (MPa√m)

Figure 7.1-2. Crack Velocity Curve for Sapphire in both Moist Air and Ringer’s Solution 
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Figure 7.1-3. Crack Velocity Curves for Sapphire Exposed to Air showing Increased Crack 

Growth Velocity with Increased RH 
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Figure 7.1-4. Comparison of the Environmental Fatigue Crack Growth Characteristics of 

Sapphire in Moist Air and Ringer’s Solution (NaCl, Na lactate, KCl, CaCl) 

The technical information, sent by the LSP to the team for review, includes pictures of two 

shattered TFAs (Figure 7.1-1) taken after removing the TFAs from their circuit cards. The 

pictures show completely disconnected shards of alumina. It seems impossible that traces 

crossing these fracture lines could have remained intact during all the stresses applied to them 

while these alumina substrates were still mounted to the circuit card. There should have been no 

problem detecting these cracks. 

Other pictures of still-mounted TFAs show almost invisible "hair line" cracks that extend for a 

limited distance along the alumina, and not all the way across a TFA. These partial cracks seem 

to have the potential to partly separate a conducting trace, but not to separate it completely. The 

resistance of the partially separated trace would increase somewhat, but not to essentially an 
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open circuit. Boeing's "enhanced" detection method is aimed (in part) at detecting the partially 

cracked traces. The completely open cracked trace should be easily detected. The other aim of 

the "enhanced" method is to detect if any TFA resistors that have been "clipped" by a crack were 

altered, but not opened electrically. 

A quantitative computation of the increase in resistance caused by a "cut" normal to the edge of a 

trace extending partway through it, shows that the increase in resistance remains essentially 

invisible until the cut is almost complete, more than 90 percent through. In cases in which there 

is "in-line" resistance from an interposed resistive element, the crack would have to extend more 

than 99+ percent through in order to be detectable as an increase in resistance by more than 5 

percent. 

The reason that the increase in trace resistance caused by a partial cut is so hard to detect is that 

the increase comes from a disturbance in the flow of the electrical current that is limited to the 

immediate neighborhood of the cut (i.e., over a length about equal to the width of the trace), 

while the total resistance comes from the entire length of the trace, and the ratio of the width to 

the total length is small. Also, in this application, the trace-resistance is not measured by itself in 

isolation, but is summed with in-line resistors, which makes trace resistance increases even more 

difficult to detect. 

Even the "enhanced" detection method, employed by Boeing to detect cracks that partially cross 

a trace, will be insensitive to partial cuts. Such a method will see almost complete cuts only. 

The Boeing "enhanced" detection scheme may or may not see cracks that pass through part of a 

film resistor. The visibility of such "clipping cuts" requires a different analysis. 

Detecting cracks in ceramics is a problem that has been around since the start of NASA’s use of 

ceramics. One of the more powerful methods is to use "vicinal illumination2

", granted the 

alumina is translucent and not dark or black, and the optical inspection is possible. 

Unfortunately, it is too late (without massive impact) to inspect the assemblies mounted in the 

Delta II about to be launched. It is safe to assume that there will be further use of sintered 

alumina in E-packaging, and the use of this inspection method may be helpful in the future. 

Figure 7.1-5 shows an image at ~50X of a slab of alumina with a pair of cracks making a Vshape. The illuminator was arranged to put a bright spot onto one side of the crack-pair. It is 

evident that the light did not cross the crack-boundaries, making these especially visible. These 

cracks were not visible in normal top-down illumination. 

2

 The vicinal illumination technique has been successfully used to detect cracks in many types of ceramic materials. 

This inspection method has become an important tool for inspection and analysis of many ceramic materials for 

space flight applications. 
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Figure 7.1-6 is the same image, but mildly enhanced to better show the crack. This image is the 

ratio of Figure 7.1-5 to an image of precisely the same spot when used to illuminate a featureless 

region of the ceramic. These two figures illustrate how powerful this non-destructive inspection 

method is. 

Figure 7.1-5. Pair of Cracks at ~50X of a Slab of Alumina (V-shaped Cracks) 

Figure 7.1-6. Ratio of Figure 7.1-5: Closer View of Crack Image
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8.0 Findings, Observations, and Recommendations 

Team findings and recommendations from this assessment are listed below. 

8.1 Findings 

F-1. A cracked TFA was detected in a Delta II first stage E-package, which had been built 

with "non-suspect" TFA. 

F-2. The Deep Impact launch vehicle has three E-packages of potential concern, one in the 

first stage and two in the second stage. The first stage E-package was manufactured in 

Monrovia, and, to date, none of these units have been found to experience cracks. The 

second stage E-packages went through ATP five times, with the enhanced screening 

applied. They were most recently tested in July 2004. 

F-3. The TFA crack risk should be considered to be in the range of medium-to-low risk, and 

represents an acceptable risk for flight. The NESC team concurred with KSC’s S&MA 

that “5x2” more accurately characterizes this risk. 

Rationale (F-1 through F-3): 

• Although the comprehensive review conducted by Boeing indicated that thick 

film cracking occurred as a result of manufacturing processing, and ATP testing 

and analysis indicated that crack propagation, which could lead to faulty TFA 

performance during flight, is unlikely, there is a finite probability that marginally 

cracked TFAs could have passed existing ATP testing criterion. 

• The LSP acknowledges the probability of failure to be “less than 1 percent”, but 

not zero, as assessed by their subjective methods. 

• The likely root cause - excessive conformal coating under the TFA assemblies 

due to manufacturing changes - has been postulated, but has not been verified by 

testing. 

• Enhanced screening test, which reviews parametric data from box level testing 

looking for changes, has proven effective to-date, but may not be 100 percent 

effective because it is an indirect method of detection. 
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• Because of the potential for undetected micro-cracks to propagate with time or 

from the launch environment, a residual risk remains for undetected micro-cracks 

to trigger failure of an assembly under stress. 

• KSC S&MA and the NESC team agree that the TFA cracking hazard represents a 

reportable risk for FRR. 

F-4. KSC S&MA, Boeing and the LSP agreed on the nature and cause of the TFA cracking 

hazard, and on the existing rationale for flight. However, when using a common 

assessment tool, they used different rationale in assigning risk codes. 

Rationale (F-4): 

• The LSP Risk Management Plan (KSC-PLN-2130) defines the risk probability 

code levels (1-5) with both verbal labels and/or numerical percent probability 

ranges. 

• KSC S&MA, Boeing and the LSP all seek to maintain internal consistency in 

their risk assessments, for continuity of reporting and internal risk management. 

• KSC S&MA does not use the percentages as guidance for code assignment, 

relying more on the verbal and subjective guidance provided in the Plan. 

• Boeing uses a subjective assessment of numerical probability as their guide to risk 

probability assignment, as provided in the LSP Risk Management Plan. 

• KSC S&MA, Boeing and the LSP apply subjective engineering judgments in 

assessing risk and assigning risk codes, as sufficient data is seldom available. 

F-5. The flight environmental loads are lower than the box ATP loads. 

Rationale (F-5): 

• Thermal cycle test margin is a minimum of +39o

 F. 

• Shock test margin is greater than 15 dB. 

• Random Vibration test margin is 5.2 dB minimum. 

8.2 Recommendations 

R-1. The Director of the OS&MA should direct the Deep Impact FRR process to continue 

with the acceptance of flight rationale for the TFA cracking issue with the acceptable 

risk, assessed as “5x2”. 
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Rationale R-1: 

• Two boxes with potentially suspect TFAs have been identified: one in the first 

stage and one in the second stage. Both have passed the enhanced parametric 

screening. 

• The first stage box (20142) was produced in Monrovia. No cracked TFAs have 

been found in Monrovia-assembled units. 

• The second stage box (20166), even though assembled in El Paso, passed the 

enhanced screening with five data sets reviewed, the most recent being July 2004. 

• Flight environmental loads are far less than the box-level ATP loads than the units 

passed successfully. 

R-2. The Director of KSC S&MA should initiate a review of the LSP Risk Management Plan 

(KSC-PLN-2130) to clarify the guidance for probability level (percent) selections, and 

make application of the tool more uniform. 

R-3. Boeing should continue to pursue other design approaches to further enhance robustness 

and minimize the likelihood of this type of failure in TFA circuits. Areas recommended 

include thicker substrates, chamfered edges, handling improvements to preclude the early 

development of micro-cracks, proof testing, and vicinal illumination inspection. 

R-4. For future missions, the Boeing and the LSP should: 

• Determine whether enhanced test(s) can be run on the vehicle, which would add 

confidence without removing the boxes. 

• Perform box-level testing and detailed data screening review of all boxes 

containing TFA, preferably within one year of launch, to guard against the 

possibility of micro-crack growth. 

• Monitor the E-packs during environmental testing to detect electrical intermittent 

conditions that may indicate a crack. 

• Consider the Destructive Physical Analysis (DPA) of an E-pack, which has 

passed the enhanced screening in 2002 but is from the “suspect” units, to 

determine whether there are signs of cracking with time. 
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9.0 Lessons Learned 

1. The only way for a complicated system such as a launch vehicle to be reliable is for each 

of its individual subsystems to be remarkably reliable. When subsystems that are less 

than reliable are accepted, failures will occur. 

2. Effective risk management requires that risks be categorized and communicated with a 

common understanding. Improved methods at consistency in accurately assessing risk 

will enhance safety and mission success. 

10.0 Definition of Terms 

Corrective Actions Changes to design processes, work instructions, workmanship practices, 

training, inspections, tests, procedures, specifications, drawings, tools, 

equipment, facilities, resources, or material that result in preventing, 

minimizing, or limiting the potential for recurrence of a problem. 

Finding A conclusion based on facts established during the assessment/inspection 

by the investigating authority. 

Lessons Learned Knowledge or understanding gained by experience. The experience may 

be positive, as in a successful test or mission, or negative, as in a mishap 

or failure. A lesson must be significant in that it has real or assumed 

impact on operations; valid in that it is factually and technically correct; 

and applicable in that it identifies a specific design, process, or decision 

that reduces or limits the potential for failures and mishaps, or reinforces a 

positive result. 

Observation A factor, event, or circumstance identified during the 

assessment/inspection that did not contribute to the problem, but if left 

uncorrected has the potential to cause a mishap, injury, or increase the 

severity should a mishap occur. 

Problem The subject of the independent technical assessment/inspection. 

Recommendation An action identified by the assessment/inspection team to correct a root 

cause or deficiency identified during the investigation. The 

recommendations may be used by the responsible C/P/P/O in the 

preparation of a corrective action plan. 
NASA Engineering and Safety Center 

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Root Cause Along a chain of events leading to a mishap or close call, the first causal 

action or failure to act that could have been controlled systemically either 

by policy/practice/procedure or individual adherence to 

policy/practice/procedure. 

11.0 Minority Report 

There were no minority opinions voiced during the performance of this consultation. 

Volume II: Appendices 

Appendix A NESC ITA/I Request Form (NESC-PR-003-FM-01)

Appendix B References 

Appendix C Cutbar Essay 

Appendix D List of Acronyms 
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Appendix A. NESC ITA/I Request Form (NESC-PR-003-FM-01) 
NASA Engineering and Safety Center 

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NASA Engineering and Safety Center 

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Appendix B. References 

1. B. Asoo, J. M. McNaney, Y. Mitamura, and R.O. Ritchie, "Cyclic Fatigue-Crack 

Propagation in Sapphire in Air and Simulated Physiological Environments", Journal of 

Biomedical Materials Research, vol. 52, 2000. 

2. Y.S. Shiue & M. J. Matthewson (2001) “Stress dependent activation entropy for dynamic 

fatigue of pristine silica optical fibers” J. Appl. Phys, 89 [9] 4787-93. 

3. NASA Presentation: “Cracked Thick Film Assemblies, Delta II, Stage 2 Servo Amplifier 

CCA EBR-02043, contributions by: Thomas Bulk et al. 

4. Leidecker, Henning, American Institute of Physics Handbook, essay entitled CutBar, 

based on a rigorous result for the increase in resistance caused by a "model crack" in a 

conductor having a rectangular cross section, page 5-24. 
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Appendix C. Cutbar Essay 
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Appendix D. List of Acronyms 

Al2O3 aluminum tri-oxide 

ATP Acceptance Test Procedure 

BATC Ball Aerospace and Technology Corporation 

CaCl Calcium Chloride 

CCAFS Cape Canaveral Air Force Station 

CF Corrosion Fatigue 

CIT California Institute of Technology 

DPA Destructive Physical Analysis 

FRR Flight Readiness Review 

GNC Guidance Navigation & Control 

GSFC Goddard Space Flight Center 

H2O Water 

ITA/I Independent Technical Assessment/Inspection 

JPL Jet Propulsion Laboratory 

JSC Johnson Space Center 

KCl Potassium Chloride 

Km kilometers 

KSC Kennedy Space Center 

LaRC Langley Research Center 

LSC Launch Services Contractor 

LSP Launch Services Program 

NaCl Sodium Chloride 

NASA National Aeronautics and Space Administration 

NCE NESC Chief Engineer 

NDE NESC Discipline Expert 

NESC NASA Engineering and Safety Center 

NRB NESC Review Board 

OS&MA Office of Safety and Mission Assurance 

RH Relative Humidity 

S&MA Safety and Mission Assurance 

SCC Stress Corrosion Cracking 

SLC Space Launch Complex 

SMARR System Mission Assurance Readiness Review 

SSP Space Shuttle Program 

TFA Thick Film Assembly 

UMD University of Maryland 
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Approval and Document Revision History 

Approved: Original signed on file 11/8/05 

NESC Director Date 

Version Description of Revision Office of Primary 

Responsibility

Effective 

Date 

1.0 Initial Release NESC Director of Safety 

Office 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Deep Impact Delta II Launch Vehicle Cracked Thick Film Coating on 

Electronic Packages Technical Consultation Report

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Cameron, Kenneth D.; Kichak, Robert A.; Piascik, Robert S.; 

Leidecker, Henning W.; Wilson, Timmy R.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-19682 NESC-RP-05-126/04-093-E

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 15-Launch Vehicles and Launch Operations

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

The Deep Impact spacecraft was launched on a Boeing Delta II rocket from Cape Canaveral Air Force Station (CCAFS) on 

January 12, 2005. Prior to the launch, the Director of the Office of Safety and Mission Assurance (OS&MA) requested the 

NASA Engineering and Safety Center (NESC) lead a team to render an independent opinion on the rationale for flight and the 

risk code assignments for the hazard of cracked Thick Film Assemblies (TFAs) in the E-packages of the Delta II launch 

vehicle for the Deep Impact Mission. The results of the evaluation are contained in this report.

15. SUBJECT TERMS

NESC; OS&MA; SCC; SSP; TFA

18. NUMBER

 OF 

 PAGES

46

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

December 2004 - November 2005

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

869021.05.07.02.99

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2009-215747

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1. REPORT DATE (DD-MM-YYYY)

01- 05 - 2009