NASA/TM-20 1 0-216293R 

· 

N~~J\ ' ! . 

Testing Requirements for Refractory 

Materials 

Luz Marina Calle 

NASA Kennedy Space Center 

Paul E. Hintze 

NASA Kennedy Space Center 

Christopher R. Parlier 

NASA Kennedy Space Center 

Jerome P. Curran 

ASRC Aerospace Corporation 

Mark R. Kolody 

ASRC Aerospace Corporation 

Jeffrey W. Sampson 

NASA Kennedy Space Center 

Eliza M. Montgomery 

NASA Postdoctoral Program 

February 2011 
NASA/TM-20 1 0-216293R 

The NASA STI Program Office ... in Profile 

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the advancement of aeronautics and space 

science. The NASA Scientific and Technical 

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NASNTM- 2010-216293R 

~ . . ' I . . ·. 

Testing Requirements for Refractory 

Materials 

Luz Marina Calle 

NASA Kennedy Space Center 

Paul E. Hintze 

NASA Kennedy Space Center 

Christopher R. Parlier 

NASA Kennedy Space Center 

Jerome P. Curran 

ASRC Aerospace Corporation 

Mark R. Kolody 

ASRC Aerospace Corporation 

Jeffrey W. Sampson 

NASA Kennedy Space Center 

Eliza M. Montgomery 

NASA Postdoctoral Program 

National Aeronautics and 

Space Administration 

Kennedy Space Center 

February 2011 
NASNTM-20 1 0-216293R 

Acknowledgments 

The research team gratefully acknowledges the contributions made by the following individuals: 

a. Nancy P. Zeitlin, Technology Integration Manager, NASNJohn F. Kennedy Space 

Center 

b. Karen Thompson, Chief Technologist, NASNJohn F. Kennedy Space Center 

c. Frank Peri, Exploration Technology Development Program Manager, 

NASNLangley Research Center 

d. Karen Whitley, Structures Materials and Mechanisms Project Manager, 

NASNLangley Research Center 

e. Judith J. Watson, Structures Materials and Mechanisms Project Manager, 

NASNLangley Research Center 

f. Scott Kenner, NASNLangley Research Center 

g. Phil Weber, Constellation Ground Operations Project, NASNJohn F. Kennedy Space 

Center 

h. Jose Perez Morales, Pad Senior Project Manager, NASNJohn F. Kennedy Space Center 

1. Dr. David Trejo, Professor, Oregon State University 

J. Cori Buchert, NASA Co-op, University of Washington 

k. Stephen Perusich, ASRC Aerospace Corp. 

l. Teddy Back, ASRC Aerospace Corp. 

m. Jolet Larracas, Structures, United Space Alliance 

n. Kevin Decker, Instrumentation Engineer, NASNJohn F. Kennedy Space Center 

o. Keith Laufenberg, Project Manager, United Space Alliance 

p. Ken Page, Ground Systems Instrumentation, United Space Alliance 

q. Kevin Taylor, Ground Systems Instrumentation, United Space Alliance 

r. WulfEckroth, Launch Site Design Manager, United Space Alliance 

s. Armand Gosselin, Launch Site Design Manager, United Space Alliance 

t. Sathish Sounderrajan, Launch Site Design United Space Alliance 

u. Gabor Tanacs, Launch Site Design Engineer, United Space Alliance 

Available from: 

NASA Center for AeroSpace Information (CASI) 

7121 Standard Drive 

Hanover, MD 21076-1320 

(30 1) 621-0390 

National Technical Information Service (NTIS) 

5285 Port Royal Road 

Springfield, VA 22161-2171 

(703) 605-6000 

This report is also available in electronic form at URL http://www.sti.nasa.gov/ and 

http://ntrs.nasa.gov/search.jsp. 

11 
Contents 

1 

1.1 

1.2 

1.3 

1.4 

2 

2.1 

2.2 

2.3 

2.4 

2.5 

2.5.1 

2.5.2 

2.6 

2.7 

2.8 

3 

3.1 

3.1.1 

3.1.2 

3.2 

3.2.1 

3.2.2 

3.2.3 

3.2.4 

3.2.5 

3.2.6 

3.2.7 

3.2.8 

3.2.9 

3.2.10 

3.2.11 

4 

4.1 

4.1.1 

4.1.2 

4.2 

5 

5.1 

5.2 

·NASA/TM-2010-216293R 

INTRODUCTION ............................................................................................. 1 

Background ........................................................................................................ 1 

Flame Deflector System ..................................................................................... 2 

Launch Environment ........................................................................................ : .6 

History of Refractory Material Testing at KSC ................................................. 8 

CURRENT REQUIREMENTS ........................................................................ 11 

KSC-GP-1059, KSC-DD-818-TR, and KSC-DM-3649 .................................. 12 

KSC-STD-Z-0004 Section 3.3.9 ...................................................................... 12 

KSC-STD-164, Section 3.3.16 ......................................................................... 12 

KSC-STD-G-0003 ........................................................................................... 13 

KSC-SPEC-P-00 12 .......................................................................................... 13 

Current Specification ....................................................................................... 13 

Revised Specification ....................................................................................... 14 

NASA-STD-5005 ............................................................................................ 16 

Product Testing ................................................................ .' ............................... 16 

Summary .......................................................................................................... 16 

LABORATORY TESTING ............................................................................. 17 

Sample Preparation .......................................................................................... 19 

Sample Preparation- Shotcreting or Gunning ... ; ............................................ 19 

Sample Preparation - Casting ............................................ :······························22 

Laboratory Testing and Standards ....................... ~ .......................................... .23 

Sieve Analysis .................................................................................................. 23 

Specific Heat Analysis ....... : ............................................................................. 24 

Cold Compressive Strength ................................................ : ........................... .24 

Abrasion Resistance ......................................................................................... 25 

Cold Modulus ofRupture ............................................................................... .26 

Thermal Expansion and Creep of Refractories ............................................... .28 

Thermal Shock ................................................. · ................................................ 28 

Thermal Conductivity ...................................................................................... 28 

Acid Resistance Tests ...................................................................................... 29 

Density, Porosity, Water Absorption, and Apparent Specific Gravity ........... .29 

Petrographic Analysis of Refractory Materials ............................................... .31 

ROCKET ENGINE EXHAUST TESTING ................................................... .31 

LC-39 Full-Scale Testing ................................................................................. 31 

Installation and Prelaunch Analysis ................................................................. 34 

Postlaunch Analysis ......................................................................................... 34 

Subscale Rocket Engine Exhaust Test Stands and Material Dimensions ........ 35 

SUBSCALE PARAMETERS .......................................................................... 39 

Launch Environment Assessment ................................................................... .39 

Parameters for Subscale T.esting ..................................................................... .43 

111 
NASA/TM-2010-216293R 

5.2.1 

5.2.2 

5.2.3 

5.3 

6 

6.1 

6.1.1 

6.1.2 

6.1.3 

6.1.4 

6.1.5 

6.1.6 

6.2 

6.2.1 

6.2.2 

6.2.3 

6.3 

7 

8 

APPENDIX A. 

APPENDIX B. 

APPENDIXC 

APPENDIX D. 

APPENDIX E. 

APPENDIX F. 

APPENDIX G. 

APPENDIX H. 

Temperature ..................................................... · ......................................... .-...... 44 

Pressure ...................................................................................................... ~ ..... 45 

Calorimeter and Radiometer ........ , .................................................................. .46 

Limitations of Subscale Testing ...................................................................... .48 

ENGINE TEST STAND FACILITIES .......................................................... .49 

John C. Stennis Space Center ................... ~ ................................ , ..................... .49 

A-1 and A-2 Test Stands ....... : ................................................... .' ...................... 49 

B-1 and B-2 Test Stands .................................................................................. 51 

E-1 Test Facility ............................................................................................... 53 

E-2 Test Facility ............................................................................................... 54 

E-3 Test Facility ............................. ' .................................................................. 55 

Building 3300 at Stennis Space Center. ........................................................... 51 

George C. Marshall Space Flight Center (MSFC) ........................................... 59 

Dynamic Test Stand ......................................................................................... 59 

Load Test Annex Facility ................................................................................. 60 

Advanced Engine Test Facility ................................................................... : .... 61 

A TK Aerospace Systems ................................................................................. 62 

LIST OF REFERENCES ................................................................................. 64 

ENDNOTES ........................................................................ : ........................... 66 

MMA-1918-80, STS-1 .............................................. , ...................................... 69 

MTS-505-81, STS-2 ............... .-........... : ........... ,. .............................................. 135 

MTS-142-82, STS-3 ................................................................................ , ..... 169 

MTS-340-82, STS-1, -2, AND 3 ············································:····················:.177 

MTS-425-82, STS-4 ...................................................................................... 193 

MTB-503-83, STS-5, -6, AND -7 .................................................................. 199 

MTB-250-84, STS-8 AND -9 ........................................................................ 215 

93-4436, STS-55 ......................................................... : ....... .' .......................... 229 

) 

IV 

, 
Figures 

Figure 1. 

Figure 2. 

Figure 3. 

Figure 4. 

Figure 5. 

Figure 6. 

Figure 7. 

Figure 8. 

Figure 9. 

Figure 10. 

Figure 11. 

Figure 12. 

Figure 13. 

Figure 14. 

Figure 15. 

Figure 16. 

Figure 17. 

Figure 18. 

Figure 19. 

Figure 20. 

Figure 21. 

Figure 22. 

Figure 23. 

Figure 24. 

Figure 25. 

Figure 26. 

Figure 27. 

Figure 28. 

Figure 29. 

Figure 30. 

Figure 31. 

Figure 32. 

Figure 33. 

Figure 34. 

Figure 35. 

Figure 36. 

Figure 37. 

Figure 38. 

Figure 39. 

Figure 40. 

Figure 41. 

NASA/TM-2010-216293R 

Launch of Apollo 11 on July 16, 1969 ............................................................. .3 

Cross Section of Flame Deflector at Launch Complex 39A. ............................ .4 

Openings for Flames From the Main Engine and SRBs ................................... .4 

Magnified View ofLC-39A Flame Deflector ............................................ , ...... 5 

SSME Flame Trench and Deflector ................................................................... 5 

SRB Main Flame Deflector ............................................................................... 7 

Side Flame Deflector ......................................................................................... 7 

History of Shuttle Launches, 1981 - 2_0 1 0 ......................................................... 9 

Flow Chart Highlighting the Developmental Path From Engineering 

Design Requirements to NASA/KSC Standards ............................................. 11 

Forms Designed To Hold the Wet Refractory Concrete (Units in Inches) ..... .20 

Position of the Forms During Spray Application ........................................... .21 

Sample Preparation (Zampell Refractories in Tampa, Florida) ...................... .21 

Example of Cut Pattern for Each Sprayed Sample (Units in Inches) .............. 22 

Paddle Mixer Used at KSC To Prepare Refractory Samples ........................... 23 

Instron in the Material Testing Laboratories, O&C Building, KSC ................ 25 

Cylindrical Sample Prior To and After Crushing ........................................... .25 

Modulus ofRupture Testing- Fondu Fyre (Pad Sprayed), Sample 81, 

Control, MOR = 837 psia (Before Compression) ............................................ 27 

Modulus of Rupture Testing- Fondu Fyre (Pad Sprayed), Sample 81, 

Control, MOR = 837 psia (After Compression) ............................................. .27 

Typical Three Sample Test Stand ................................................................... .32 

Test Stand Locations ........................................................................................ 32 

Typical Sample Size From KSC-SPEC-P-0012 .............................................. 33 

.Pretiminary Test Stand Design, Top View ...................................................... 36 

Cross-Section Details From Top View ........................................................... .37 

Test Stand Side View ....................................................................................... 37 

Refractory ........................................................................................................ 38 

Refractory Material Sample Cross-Sections .................................................... 38 

Sensor Locations on the East Wall of the SRB Flame Trench ....................... .40 

Sensor Locations on the West Wall of the SRB Flame Trench ...................... .41 

Location of Pressure Sensor 9W Underneath the MLP .................................. .41 

Sensor Locations on the SRB and SSME Flame Deflectors ........................... .42 

Location 4E Is Highlighteq in the Circle ........................................... : ............ .43 

Selected Temperature Measurements During the Launch of STS-125 .......... .45 

Selected Pressure Measurements During STS-126 ......................................... .46 

Selected Calorimeter Measurements During STS-119 ................................... .47 

Selected Radiometer Measurements During STS-126 ................................... .48 

Exterior View of the A-2 Test Stand ............................................................... 50 

View of the A-2 Flame Duct and Diffuser ....................................................... 51 

Rocket Motor Test at Stennis B-1 Complex .................................................... 52 

Stennis B-2 Flame Deflector ............................................................................ 53 

Stennis E-1 Flame Deflector ............................................................................ 53 

E-2 Cell 1 Test Stand at Stennis Space Center ................................................ 54 

v 

') 
NASA/TM-2010.,.216293R 

Figure 42. 

Figure 43. 

Figure 44. 

Figure 45. 

Figure 46. 

Figure 47. 

Figure 48. 

Figure49. 

Figure 50. 

Figure 51. 

Figure 52.· 

Figure 53. 

Figure 54. 

Tables 

Table 1. 

, Table 2. 

Table 3. 

·Table 4. 

Table 5. 

Table 6. 

Table 7. 

Table 8. 

· Stennis Space Center E-2 2 Vertical Test Flame Duct .................................... 55 

E-3 Vertical Configuration Test Cell Flame Duct.. ......................................... 56 . . 

E-:3 horizontal Test Configuration ................................................................... 51 . 

Test Duct at Stennis Space Center ................ :;'; ................................................ 51 

Refractory Concrete Apron at Stennis Space Center .......... _ ........ : .................... 58 

Stennis Plume Deflector Test Rig ....... : ....... : ............... ~ ..•................................. 59 

MSFC ,Dynamic Test Stand : ......................................................... ; .................... 60 

MSFC Test Stand 116, Solid Fuel Torch Test.. ....... : ......... ~ .................. : .......... 61 

MSFC Test Stand 115, Solid Fuel Torch Test.. ... : ........................................... 61 

A Scaled-Down 24-Inch Version of the Space Shuttle's Reusable Solid 

Rocket Motor Fired at a MSFC Test Stand ..................................................... 62 

Horizontal Testing of an ASAS 21-120 Motor. ............................................... 63 

Space Shuttle RSRM ....... : ........... ~ .......................................................... : ......... 63 

Space Shuttle RSRM Test Firing at the ATK, Promontory, Utah Facility ...... 64 . 

Overview of Test Program Since 1981 ....................................... ; .................... 10 

Laboratory Evaluation Requirements for Refractory Concrete, Sugg~sted 

New Revisio~ to KSC-SPEC-P-0012 ................................. : ............... ; ............ 15 · 

Summary of Testing and Qualification Requirements for Refractory 

Materials in Launch Environments .................. : ................................... : ........... 17 

Sensors Used During Launch Environment Assessment.. ............................. ..40 · 

Maximum Temperature, Pressure,' and Heat-Flux Measured During the 

Launch ofSTS-126, STS-119, STS-125, and STS-127 ..................... ~ ............ .44 

Maximum Temperatures (°F) Measured During Three Launches .................. .44 

Maximum and Minimum Pressures (psi g) Measured During Three . i 

Launches ........................................................................................ · .................. 46 

Maximum Calorimeter Measurements (btu/ft2 sec) Obtained During Two 

Launches ................................................................... ~ ...................................... 47 

Vl 
ACI 

AIAA 

AISC 

Alz03 

ANSI 

ASCE 

ASRB 

ASRM 

ASTM 

Btu 

CAC 

COTS 

CP 

CxP 

DSC 

DTS 

EAR 

ECO 

EDS 

ETDP 

FF 

FOD 

g 

GH2 

GHe 

GN2 

GO 

gpm 

GSE 

HB 

HCl 

ABBREVIATIONS AND ACRONYMS 

American Concrete Institute 

American Institute of Aeronautics and Astronautics 

American Institute of Steel Construction 

aluminum oxide 

American National Standards Institute 

American Society of Civil Engineers 

Advanced Solid Rocket Booster 

Advanced Solid Rocket Motor 

American Society for Testing and Materials 

British thermal unit 

calcium aluminate cement 

commercial off the shelf 

center perforated 

Constellation program 

differential scanning calorimetry 

Diagnostic Test Facility 

Export Administration Regulation 

Export Control Office 

energy dispersive x-ray spectroscopy 

·Exploration Technology Development Program 

FonduFyre 

foreign-object-debris 

gram 

gaseous hydrogen 

gaseous helium 

gaseous nitrogen 

ground operations 

gallon per minute 

ground support equipment 

NASA Handbook 

hydrochloric acid 

Vll 

NASA/TM-2010-216293R 
NASA/TM-2010-216293R 

IBC 

ITAR 

KSC 

LC 

LHz 

L02 

LOX 

MFD 

MIL-STD 

MLP 

MOR 

MPa 

MSFC 

MSL 

NASA 

NLS 

OSHA 

ps1 

psi a 

RSRM 

s 

SEM 

SFD 

SLC 

SMM 

SPEC 

SPREE 

SRB 

SRM 

SSe 

SSME 

STD 

STI 

International Building Code 

International Traffic in Arms Regulations 

Kennedy Space Center 

Launch Complex 

liquid hydrogen 

liquid oxygen 

liquid oxygen 

main flame deflector 

military standard 

Mobile Launcher Platform 

modulus of rupture 

megapascal 

· George C. Marshall Space Flight Center 

Materials Sciences Laboratory 

National Aeronautics and Space Administration 

National Launch System 

Occupational Safety and Health Administration 

pound per square inch 

pound per square inch absolute 

reusable solid rocket motor 

second 

scanning electron microscopy 

side flame deflector 

Space Launch Complex 

Structures, Mechanisms, and Materials 

specification 

solid propellant rocket exhaust effects 

solid rocket booster 

solid rocket motor 

Stennis Space Center 

Space Shuttle Main Engine 

standard 

Scientific and Technical Information 

Vlll 
NASA/TM-2010-216293R 

STS Space Transportation System 

TBD to be determined 

TM Technical Memorandum 

TPP Technology Prioritization Panel 

USA United Space Alliance 

IX 
NASA/TM-2010-216293R 

This page intentionally left blank. 

X 
NASA/TM-2010-216293R 

ABSTRACT 

Launch Pads 39A and 39B currently use refractory material (Fondu Fyre) in the flall)e trenches. 

This material was initially approved for the Saturn program, had a lifetime of 10 years according 

to the manufacturer, and has been used for over 40 years. As a consequence, the Fondu Fyre at 

Launch Complex 39 requires repair subsequent to almost every launch. 

With the recent severe damage to the flame trenches, a new refractory material is sought to 

replace Fondu Fyre. In order to replace Fondu Fyre, a methodology to test and evaluate 

refractory products was developed. This paper outlines this methodology and discusses current 

testing requirements, as well as the laboratory testing that might be required. Furthermore, this 

report points out the necessity for subscale testing, the locations where this testing can .be 

performed, and the parameters that will be necessary to qualify a product. The goal is to identify 

a more durable refractory material that has physical, chemical, and thermal properties suitable to 

withstand the harsh environment of the launch pads at KSC. 

Xl 
NASA/TM-2010-216293R 

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xu 
NASA/TM-2010-216293R 

TESTING REQUIREMENTS FOR REFRACTORY MATERIALS 

1 INTRODUCTION 

1.1 Background 

Corrosion is the environmentally-induced degradation of materials. The natural marine 

environment at the Kennedy Space Center (KSC) has been documented by the American Society 

for Metals (ASM) as having the highest corrosion rate of any site in the continental United 

States. See Endnote 1 in Section 8. As a result, launch structures and ground support equipment 

(GSE) at KSC degrade faster than similar assets at other locations. With the introduction of the 

Space Shuttle in 1981, the already highly corrosive natural conditions at the launch pads were 

rendered even more severe by the acidic exhaust from the solid rocket boosters (SRBs ). As a 

consequence, corrosion-related costs are significant for all launch structures. These costs were 

estimated,in Januaiy 2009 to be approximately $336M over the previous 20 years of the Space 

Shuttle Program. The estimate inCluded the costs associated with inspection and maintenance of 

the launch pads, medium-scale and large-scale blasting and repainting activities, the repair and 

replacement of failed refractory materials, and the replacement ofbadly corroded structural 

metal elements. Technologies for the prevention, detection, and mitigation of materials 

degrad~tion in launch facilities and ground support equipment were identified by the 

Constellation Program Ground Operations (CxP GO) as a critical need for the safety, efficiency, 

affordability, and sustainability of future launch operations at KSC. Subsequently, CxP GO 

established an agreement with the Exploration Technology Development Program (ETDP) 

Structures, Mechanisms, and Materials (SMM) project to identify alternate refractory material 

for the protection of the launch pad flame deflectors at KSC. This report, prepared as one of the 

deliverables for the project, provides testing requirements for refractory materials to be used in 

launch pad applications as well as the available information from all previous testing of 

refractory materials for launch pad flame trench protection that was gathered in the process of 

identifying the testing requirements 

Materials development, characterization, testing, and optimization would be based on 

identification of key refractory material performance properties that can improve the durability, 

performance, and safety of the launch complexes. A small-scale prototype flame deflector 

system would be developed (or acquired) for component-level materials testing. Material 

requirements; maintenance and inspection requirements; application, repair, and rehabilitation 

requirements; system specifications; and qualification requirements/procedures would be 

developed for the replacement refractory material. These refractory material technologies were to 

be integrated into a scaled, simulated flame deflector system and demonstrated under simulated 

launch conditions. 

The above project plan had to be revised because of the unavailability of funding to carry out the 

project as it was originally planned. This report constitutes the last deliverable for the scaleddown project, and it is intended to provide the testing requirements for refractory materials to be 

used in launch pad applications, as well as provide the available information on all previous 

testing of refractory materials for launch pad flame trench proteCtion that was gathered in the 

process of identifying the testing requirements. 

1 
NASA/TM-2010-216293R 

1.2 Flame Deflector System 

The launch complexes at the KSC are critical support facilities required for the safe and 

successful launch of vehicles into space. Most of these facilities are over 30. years old and are 

experiencing deterioration. See Figure 1. With constant heat/blast effects and environmental 

exposure, the refractory materials currently used in the NASA .launch pad flame deflectors have 

become very susceptible to failure, resulting in large pieces of refractory materials breaking 

away from the steel base structure and being projected at high speeds during launch. Repair of 

these failures is a costly and time-consuming process. Improved materials and systems for use in 

launch pad flame deflectors will improve supportability in KSC launch facilities by reducing 

operational life-cycle costs. 

The flame deflecto~ systems at Launch Complex (LC) 39A a~d I;C-39B are critical to protect 

NASA's assets, including the Space Shuttle, GSE, and personnel. As the name implies, the 

system diverts rocket exhaust away from critical structures through its geom~tric design. Further 

benefits are.provided by a water deluge system that dampens acoustic vibrations and the high 

temperatures associated with launch. 

Flame deflectors are typically covered with a heat-resistant material that protects the flame 

deflector from erosion, ablation, and extreme temperatures that are produced by the rocket 

propulsion systems. If this refractory layer is compromised, deterioration to the flame deflector 

and other load-bearing structures may result. Once compromised, the refractory material and 

flame deflector substructures can tum into unwanted projectiles known as foreign-object-debris 

(FOD) that can cause consequent damage. 

LC-39A and 39B were originally designed to support the Apollo Program and the Saturn V 

rocket. With the advent of the Shuttle Program, the Saturn-era flame deflectors were replaced. 

Figure 2 shows a schematic cross section of the current flame deflector at launch complex 39A. 

The flame deflector system consists of a flame trench, a main flame deflector (MFD), and a pair 

of side flame deflectors (SFDs ). The main flame deflector is designed in an in inverted, 

V -shaped configuration, is constructed from structural sted, and is covered with refractory 

concrete material. One side of the inverted "V". deflects the flames and exhaust from the Space 

Shuttle Main Engine (SSME), and the opposite side deflects the flames and exhaust from the 

SRBs .. Additional protection is provided by the two movable side deflectors at the top of the 

trench (not shown in the figure). The SFDs direct the SRB exhaust and are needed because the 

SRBs.are very close to the sidewalls of the flame trench. The orbiter side of the flame deflector 

is 38-feet high, 72-feet long and57-:feet wide. The SRB side of the flame deflector is 42-feet 

high, 42-feet long and 57-feet wide. The total mass of the asset is over 1 million pounds. See 

Endnote 2 in Section 8. 

2 
NASA/TM-2010-216293R 

Figure 1. Launch of Apollo 11 on July 16, 1969 

The flames from the SSMEs and the SRBs are channeled down opposite sides of the flame 

deflector. See Endnote 3 in Section 8. The deflector is constructed of steel on a structural steel 

I-beam framework. To protect the structure from serious degradation during launch, the faces of 

the flame deflector are lined with refractory concrete. This product is known as Fondu Fyre 

WA-lG (supplied by the Pryor Giggey Co.). The thickness of the refractory concrete is 

approximately 6 inches on the SRB side, 4.5 inches on the SSME side, and 4 inches on the side 

deflectors. 

3 
NASA/TM-2010-216293R 

Flames from the solid rocket 

boosters 

North Side 

Flames from orbiter's 

main engines 

Figure 2. Cross Section of Flame Deflector at Launch Complex 39A 

Figure 3 shows the configuration of the Space Shuttle viewed upward from the flame trench. The 

openings for the SSME exhaust and the flame deflector used to divert the rocket plume from the 

SRBs are labeled. The other side of the flame deflector, which is not visible in the picture, 

diverts the exhaust from the main engines. The SRBs burn at a much higher temperature and 

create a harsher loading environment than the SSME. Consequently, the SRB exhaust leads to 

more severe exposure conditions and results in damage that is more significant to the deflector. 

Figure 3. Openings for Flames From the Main Engine and SRBs 

Figure 4 shows a view of the flame deflector underneath the SRBs. The image shows the 

structural steel at the bottom of the deflector, which is protected with Fondu Fyre. Figure 5 

shows the SSME flame trench and deflector. 

4 
NASA/TM-2010-216293R 

Figure 4. Magnified View of LC-39A Flame Deflector 

Figure 5. SSME Flame Trench and Deflector 

5 
NASNTM-2010-216293R 

Safely meeting the flame deflector requirements of diverting the flame, exhaust, and small items 

that are dislodged during launch is dependent on the integrity and performance of the materials 

used to construct the flame deflectors. The use of refractory products that have superior material 

characteristics (under launch conditions) is necessary to protect the flame deflector, Space 

Shuttle, GSE, and .launch personnel. 

1.3 Launch Environment 

The launch environment is different in the SRB and SSME flame trenches. The SRB side has 

historically seen more damage than the SSME side because of the harsher conditions found 

there. This section gives a general overview of the launch environment. 

The Space Shuttle has two SRBs, which exhaust in the north flame trench, and three SSMEs, 

which exhaust towards the south. The SRBs have considerably more thrust, 3,300,000 pounds 

each, compared to the thrust of the SSMEs, 375,000 pounds each. The SRBs also bum hotter 

than the SSMEs and produce aluminum oxide particles that can act as abrasives or, if they are 

near or above melting point, may react with the refractory material. The SRBs impinge in two 

locations on the top of the flame deflector, underneath the MLP exhaust holes as seen in Figure 

6. The light areas at the top left and right are the direct impingement areas. The areas that receive 

direct impingement appear lighter because of the presence of aluminum oxide particles in these 

locations. There are two side flame deflectors above the flame trench, shown in Figure 7. The 

SRBs impinge on the side deflectors before entering the main flame deflector. Examination of 

the impingement area shows that the material sees very different conditions than outside the 

impingement area. These differing conditions may cause different failure mechanisms for the 

refractory material. For example, .the bottom lip of the deflector appears to undergo more erosion 

than those areas farther up the deflector towards its apex. 

The launch sequence itself affects the environment. Prior to launch, water is continuously flowed 

onto the refractory material. This procedural requirement ensures that the sound suppression 

system is operational and results in the refractory material being thoroughly saturated with water 

during launch. This process releases approximately 300,000 gallons of water during launch, with 

a peak flow rate of 900,000 gallons per minute 9 seconds after launch. See Endnote 4 in Section 

8. The launch timeline is as follows: 

a. The sound suppression water flow starts just before SSME ignition at T - 6.6 

seconds (s). 

b. SSME ignition occurs at T - 6.6 s. 

c. SRB ignition occurs at T - 0 s. 

d. The Shuttle clears the tower approximately 6 seconds after launch. See Endnote 5 

in Section 8. 

6 
NASA/TM-2010-216293R 

Figure 6. SRB Main Flame Deflector 

Figure 7. Side Flame Deflector 

7 
NASNTM-2010-216293R 

1.4 History of Refractory Material Testing at KSC 

Investigations on refractory materials for protection from heat and blast on launch complexes 

using solid rocket motors (SRMs) started in the mid 60's. A Solid-Propellant Rocket Exhaust 

Effects (SPREE) program (see Endnote 6 in Section 8) was established to evaluate materials for 

the protection of launch facilities from potentially damaging environments created by solid 

rocket motor exhaust gases. · · · . 

The program consisted of performing scale-model test firings of both· cold and hot jet streams 

with single nozzle configurations to derive models. Subscale test modeling was refined and 

correlated with full-scale launches of the Titan IIIC rocket using 120-inch SRMs. Unfortunately, 

this program's goal was to develop a design handbook for properly configuring a deflector 

system for launching wi~h SRMs and did not concentrate on qualifying refractory materials. 

Out of the 27 test firings performed, only five refractory materials from tw() manufacturers were 

used. Consequently, F ondu Fyre W A -1. was recommended for use because of its superior 

performance in mechanical, physical, and chemical properties over the other products. 

Furthermore, it was already approved and in use on Complex 34 for the Apollo Program. 

Since then, investigations on thirteen different refractory materials have been performed at KSC 

during the launches ofSTS-lthrough STS-9 (1981- 1983) and again on STS-55 (1993). Over 

the past 30 years (Figure 8), only 10 launches out of 131 (less than 8%) were utilized for testing 

new refractory materials. Some of the new products showed promise, but further investigations 

were not performed. These investigations mostly consisted of exposing test panels from different 

refractory concrete maimfacturers to actual launch environments (i.e., panels were placed in the 

flame deflector during a launch) at LC-39 with the objective of qualifying the materials for use in 

the flame deflectors. There is no evidence that the material used today, W A-1 G (gunnable) was 

ever tested. 

8 
NASA/TM-20 1 0-216293R 

8 ~-------,~----------------------------------------------------. 

<ll 

~ 

7 

6 

'5 5 

= 

= ~ 

....:l 

eo.. 4 

0 

I. 

~ 

,.Q 

e 3 

= z 

2 

Challenger 

Tragedy 

0 -~-&-.....-......-.......-.......-. ...... 

Figure 8. 

Year 

Columbia 

Tragedy 

History of Shuttle Launches, 1981 - 2010 

The main issue with qualifying a new material was the uncertainty of a product's compatibility 

with the existing installed system from different manufacturers. Refractory concrete panels were 

evaluated for average depth of erosion and pitting, where pitting is defined as the maximum 

depth of abrasion. Table 1 shows an overview of the test program including the number of 

samples tested and the products used for each test. Copies of each test report are provided in 

Appendix A, MMA-1918-80, STS-1 ; Appendix B, MTS-505-80, STS-2; Appendix C, MTS-142-

82, STS-3; Appendix D, MTS-340-82, STS-1 , -2, and -3; Appendix E, MTS-425-82, STS-4; 

Appendix F, MTB-503-83, STS-5, -6, and -7; Appendix G, MTB-250-84, STS-8 and -9; and 

Appendix H, 93-4436, STS-55 of this document. 

9 
NASNTM-2010-216293R 

Table 1. Overview of Test Program Since 1981 

Report No. Launch Report Date Materials Tested (#of samples) 

Fondu Fyre WAl (1)1 

Fondu Fyre WAl w/wire (1) 1 

MMA-1918-80 STS-1 July 29, 1981 WRP1(1) 2 

(Appendix A) WRP 3 (1) 2 

Tufshot (1) 3 

Tufshot w/ wire (1) 3 

Fondu Fyre WAl (1) 1 

Fondu Fyre WAl w/wire (1) 1 

MTS-505-81 STS-2 March 1, 1982 WRP 2 (1) 2 

(Appendix B) 

WRP 3 (1) 2 

Tufshot w/ wire (1) 3 

Fondu Fyre FSC-5 (1) 1 

Fondu Fyre W A-1 1 

MTS-142-82 STS-3 May 6,1982 

Fondu Fyre WA-1 w/wire (1) 1 

(Appendix C) WRP 1 w/wire (1) 2 

WRP 3 w/wire (1) 2 

Tufshot w/wire (1) 3 

Fondu Fyre WA-1 Panel2 

Fondu Fyre WA-1 w/wire Panel3 

WRP-1 Panell 

MTS-340-82 STS-1, 2, and 3 June 7, 1982 WRP-2, Panel 7 (Appendix D) 

WRP-3 Panel 4 

Tufshot Panel 6 

Tufshot with fibers, Panel 5 

MTS-425-82 STS-4 August 27, 1982 

Fondu Fyre W A-1 (3) 1 

(Appendix E) WRP Ql_2 

Fondu Fyre W A-1 (3) 1 

WRP1(1) 2 

MTB-503-83 STS-5, -6, and -7 September 1, 1983 

WRP 2 (1) 2 

(Appendix F) WRP 3 w/wire (10) 2 

LI (1) 3 

17-67 (2) 3 

Fondu Fyre HT-1 (1) 1 

Fondu Fyre FSC-5 (1) 1 

MTB-250-84 · STS-8 and -9 May 22,1984 WRP 3 w/wire (2) 2 

(Appendix G) 

17-67 (1) 3 

No. 75 _(_2) 4 

93-4436 STS-55 July 19, 1993 

Fondu Fyre WA-1 (3) 1 

(Appendix H) Mitec (3) 5 

1

0riginally Designed Concretes, now Pryor Giggey Co, 2Wahl Refractory Products, 3Harbison Walker Refractories, 4

Sauereisen Cement Co, 5Mitec, Inc 

10 
NASA/TM-2010-216293R 

2 CURRENT REQUIREMENTS 

NASA KSC design requirements dictate specific standards for selecting and qualifying 

refractory materials for use in the launch environment. A flow chart was developed to visually 

display the NASA/KSC standards that must be met under KSC-DE-512-SM, Facility, System, 

and Equipment General Design Requirements (Figure 9). This shows the design requirement 

pathway that leads to understanding and predicting the launch-induced environment and 

eventually to the standards used to meet those environment and design requirements. A 

description of the launch-induced environment and corresponding requirements, as listed in the 

documents and standards below and specifically related to refractory materials, follows. 

NS 1S OnOO Volume 

SW.£-0002 forGSE NH8 7320 for Facility 

LOADS/ENVIRONMENT 

SAFETY FACTORS/ AllDWABLES 

Figure 9. 

~ 

~ 

1 ...... !,,._,; 

KSC-STD-1648 

KSC-STO.G-00038 

KSC-SPEC-P-0012 

NASA-5TD-5005 

0 Appllcoble flnny 

Flow Chart Highlighting the Developmental Path From Engineering 

Design Requirements to NASA/KSC Standards 

11 
NASNTM-2010-216293R 

2.1 KSC-GP-1059, KSC-DD-818-TR, and KSC-DM-3649 

The documents KSC-GP-1059, KSC-DD-818-TR, and KSC-DM-3649 all describe the launchinduced environment. No testing and qualification requirements are listed in these documents but 

rather they are the supporting documents that describe launch-induced environment conditions 

for which the refractory concrete-related specifications are based. 

KSC-GP-1059, Environment and Test Specification Levels Ground Support Equipment for 

Space Shuttle System Launch Complex 39 (Acoustic and Vibration, Volume I of II; Thermal and 

Pressure, Volume II of II, Part 1 of 2, Heating Rates, Pressure Loads, and Plume Flow Fields, 

and Part 2 of2, Specific Temperatures of Selected Parts ofLC-39; Acoustics, Volume III; and 

Thermal and Pressure, Volume IV) encompasses the thermal, pressure, acoustic, and vibration 

environments induced by launch of Space Shuttle vehicles from LC-39 Pads A and B. Volumes I 

and II predict the environments from Saturn V launches and anticipated environments for Space 

Shuttle launches. Volumes III and IV predict the launch-induced environments of Space Shuttle 

launches after the replacement of Redesigned Solid Rocket Motors (RSRM) and Solid Rocket 

Boosters (SRB) with Advanced Solid Rocket Boosters (ASRB). Volumes III and IV used 

launches from STS-1 to STS-30R to predict the new launch environments. 

KSC-DD-818-TR, SummaryofMeasurements ofKSC Launch-Induced Environmental Effects 

(STS-1 through STS-11), summarizes the Shuttle launch-induced environment data acquired at 

Kennedy Space Center during the STS launches. The measurements included using sensors to 

record pressure, acoustic, strain, load, temperature, heat rate, and vibration parameters. 

KSC-DM-3649, Lift-off Response Spectra to Space Shuttle Launch-Induced Acoustic Pressures, 

summarizes computations of lift-off response spectra obtained from the acoustic pressure created 

and correspondingly measured during Shuttle launches. This document uses the acoustic data 

from STS-2 through STS-31 to derive static and dynamic loads imposed on exposed GSE during 

launching. 

2.2 KSC-STD-Z-0004 Section 3.3.9 

KSC-STD-20004, Structural Design, Standard for, defines requirements for framing of 

structures, ground support equipment, and temporary structures and enclosures. In se'ction 3.3.9, 

the standard refers to KSC-SPEC-P-0012, Refractory Concrete, Specification for, qualification 

of materials used in structures that are subject to direct rocket engine exhaust impingement. 

2.3 KSC-STD-164, Section 3.3.16 

KSC-STD-164, Environmental Test Methods for Ground Support Equipment, Standard for, 

focuses on test methods for ground support equipment, with section 3.3.16 specifically related to 

Lift-Off Blast. Refractory materials at the launch pad must meet this standard. According to this 

standard, the refractory materials can only be qualified by meeting lift-off blast test conditions. 

No other qualification methods, such as laboratory or test chamber methods, can be substituted~ 

The lift-off blast test can only be performed during an actual launch or on a rocket engine test 

stand during a test firing. 

12 
NASNTM-2010-216293R 

2.4 KSC-STD-G-0003 

KSC-STD-G-0003, Launch Support and Facility Components, Qualification of, Standard for, 

establishes methods for the qualification of launch support and facility components. Section 4.0 

describes five methods for qualifying a material: (1) qualification by testing, (2) qualification by 

similarity, (3) prior qualification, (4) qualification by usage and analysis, and (5) qualification by 

higher level assembly testing. Refractory materials must meet this standard (KSC-STD-G-0003). 

Qualification by testing includes demonstration of operational suitability for a specific 

application when no qualification date is available. Three types of tests (functional, acceptance, 

and verification) are used in the design and development of new equipment. This testing includes 

using structural, dynamic, component compatibility and life cycle criteria that is similar to the 

launch environment but on a laboratory scale. The testing allows the use of subscale accelerated 

conditions of environment, functions (such as pressure, voltage, flow, etc.), tolerances, life 

cycles, and time. This testing includes using all natural environmental factors, including 

humidity, salt fog, rain, sand and dust, fungus, and solar radiation, and then inducing launch 

environments, including acoustics, shock, vibration, high and low temperatures, liftoff blast, 

electromagnetic field, and explosive gas/vapor atmosphere. The environmental testing must be in 

accordance to KSC-STD-164, which is testing during an actual launch or rocket engine test 

firing. 

Qualification by similarity refers to materials that are comparable in use or rating to alreadyqualified materials. · 

Prior qualification refers to materials that were formerly qualified to the necessary environment 

and testing levels. · 

Qualification by usage and analysis refers to acceptance without a formal qualification test and 

based on evaluations of tis age on previous programs, postlaunch data and inspection, and 

considering all static and dynamic operating conditions. 

Qualification by higher level assembly testing refers to qualification of a component that is part 

of a higher assembly that has already been qualified by testing in the same application. 

2.5 KSC-SPEC-P-0012 

KSC-SPEC-P-0012, Refractory Concrete, Specification for, covers requirements for refractory 

concrete used for the heat and blast protection of the flame deflectors and other areas of a launch 

or test facility. The material must resist degradation of thermal protection characteristics due to 

the unprotected seacoast atmosphere exposure at the launch facilities. The KSC specification 

currently active for refractory concrete is KSC-SPEC-P-0012; however, it is currently being 

revised. 

2.5.1 Current Specification 

The current requirements for KSC-SPEC-P-0012 include a qualification process, required 

material characteristics, minimum fresh and hardened material requirements, as well as quality 

assurance provisions and packing requirements. The quality assurance section provides 

13 
NASAJTM-2010-216293R 

' information on the material qualification process for use at KSC. This process includes making 

test specimens, as required by the specification, and exposing these specimens to an actual 

launch environment. Requirements listed in the specification for the acceptance of refractoryconcrete materials used at KSC include the following: 

a. Materials: Aggregate shall be hard, dense, durable, clean, sharp, and well graded. 

b. Fineness modulus: The fineness modulus shall be between 3.75 and 2.75. 

c. Strength: The refractory concrete shall develop a compressive strength of 4500 

psi at 7 days and 90 percent of the 7 -day strength within 24 hours. 

d. Rocket engine exhaust resistance: Test samples installed at designate areas of the 

launch facility shall not crack, spall, or erode more that 1/8 inch when subject to 

rocket exhaust test. Heat flux shall be up to 3300 Btu/ft2 -sec with an exposure 

time of approximately 10 seconds. 

e. Workability: The refractory concrete shall be able to be applied pneumatically of 

manually to a smooth finish 

f. Weathering: The refractory concrete shall resist degradation of thermal protection 

characteristics due to seacoast atmosphere exposure 

Some of the limitations that have been identified for the existing KSC-SPEC-P-0012 

specifications include the following: 

a. Key performance parameters, such as the material shrinkage, are not required in 

the current specification. 

b. No requirements on material storage are provided. Calcium aluminate cement 

(CAC) can hydrate with time under storage conditions that can significantly 

change the performance of the material. 

c. · No requirements are provided on placement procedures, curing, or other key 

construction practices. Procedures for placement of the entire refractory lining are 

dependent on the material being used for the refractory lining. 

d. Nq methodology is provided for qualifying materials. Specific requirements, 

based on existing or newly developed standardized tests, should be included. 

This specification is currently being revised to include the newly-identified requirements for 

refractory concrete. 

2.5.2 Revised Specification 

The revised specification will include many of the previous requirements, as well as new 

requirements based on the above noted limitations. 

14 
NASA/TM-2010-216293R 

Table 2 lists the laboratory evaluation parameters that each material must meet in order to be 

considered for the second stage evaluation, a rocket engine exhaust exposure. The same 

laboratory evaluation requirements must then be met for the materials after exposure to the 

rocket engine exhaust. 

Table 2. Laboratory Evaluation Requirements for Refractory Concrete, Suggested 

New Revision to KSC-SPEC-P-0012 

Material Test Standard Test Requirement Characteristic 

ASTM C92, using 

Sieve analysis particle size Reportable 

distribution method 

Compressive ASTM C 133, using 4500 psi at 7 days, with 90 percent of 

strength cold crushing strength 7 -day strength developed within 24 

method hours 

No explosive spall in less than 6 thermal 

Thermal shock ASTM C1171 or C24 cycles when shocked to a temperature 

determined by the Government 

Abrasion resistance ASTMC704 Erosion not more than 3.2 mm per cycle 

Shrinkage/thermal ASTM C 832-00 TBD expansiOn 

ASTM C133, using 

Modulus of rupture modulus of rupture TBD 

method 

Acid resistance test Internal Government TBD test* 

Thermal ASTMC1113 Reportable conductivity 

ASTM E1269 from 

Specific Heat room temperature to - Reportable maxtmum use 

temperature 

Bulk Density ASTM C20, section Reportable 1.1.4 

*The acid resistance test was developed by the Government. Details of the test will be given in 

the specification and are included in this document. 

15 
NASNTM-2010-216293R 

Rocket Engine Exhaust Exposure: 

a. The test sample must be installed at a designated test or launch site and be 

exposed to engine exhaust. The test will be performed by the Government. 

b. . The test sample must meet the same testing requirements as above after the rocket 

engine exposure period. 

Additional Requirements: 

a. The material may not contain asbestos. 

b. The refractory shall be capable of meeting the requirements for a number of 

thermal cycles or number of year's exposure to the seacoast environment. · 

c. The installation, curing process, and temperatures for each phase of testing and 

final installation at the launch or test site shall be the· same. 

The reader is urged to corisult with the most current version ofKSC-SPEC-P-0012 for a 

complete list of the current and revised requirements. 

2.6 NASA-STD-5005 

NASA-STD-5005, Standard of the Design and Support of Ground Equipment, defines top-level 

requirements and provides guidance for the design and fabrication of ground support equipment. 

In section 5.11.3.1.5.3, the standard refers to KSC-SPEC-P-0012 for qualification of refractory · 

materials used in structures that are subject to heat and blast protection of flame deflectors. 

2.7 Product Testing 

In order for a product to be considered for rocket engine exhaust testing,. the product needs to 

meet the above referenced requirements. The product must be prepared, installed, and cured 

using the same methodology as intended for installation at the launch pad. 

2.8 Summary 

All of the standards and specifications regarded the exposure to la~ch environments as a critical 

step for the qualification of refractory materials. A summary of testing requirements referenced 

in different standards and specificationsforrefractory materials used in launch environments is 

shown in Table 3. This information is based on the requirements included in KSC-STD-Z0004F, KSC-STD-164, KSC-STD-G-0003; and KSC-SPEC-P-0012. 

16 
NASA/TM-2010-216293R 

Table 3. Summary of Testing and Qualification Requirements for Refractory 

Materials in Launch Environments 

NASA Qualification R~uirements 

Specification Similar Laboratory Subscale Launch Industry Scale 

KSC-STD-Z-0004F Refers to KSC-SPEC-P-0012 for requirements. 

KSC-STD-164 No No Yes Yes 

Yes, along 

KSC-STD-G-0003 Yes with KSC- Yes Yes STD-164 

requirements 

KSC-SPEC-P-0012 No Yes No Yes 

Revised No Yes Optional Yes KSC-SPEC-P-0012 

NASA-STD-5005 Refers to KSC-SPEC-P-:0012 for requirements. 

3 LABORATORY TESTING 

Standardized testing of refractory materials is required to assess the characteristics of refractory 

products and to identify products that can be used in KSC launch environments. For refractory 

materials, the following standards are applicable. It should be noted that some standards are not 

specific to refractory products, and modifications to the testing procedures may be required. In 

most cases, and when possible, the standard that specifically pertains to refractory products 

should be used. 

The following standards are applicable to the testing of refractory materials for NASA. It should 

be stressed that the list is not all inclusive of the standards that might be required for future needs 

and requirements. 

a. Preparation of Refractory Samples 

(1) ASTM C1140, Standard Practice for Preparing and Testing Specimens 

From Shotcrete Test Panels 

(2) ASTM C862, Standard Practice for Preparing Refractory Concrete 

Specimens by Casting 

b. Compressive Strength 

(1) ASTM C133 (Reapproved 2003), Standard Test Methods for Cold 

Crushing Strength and Modulus of Rupture of Refractories 

17 
NASA/TM-2010-216293R 

c. Abrasion/Erosion Resistance 

(1) ~ ASTM C704, Standard Test Method for Abrasion Resistance of 

Refractory Materials at Room Temperature 

d. Shrinkage and Thermal Expansion 

(1) ASTM C179, Standard Test Method for Drying and FiringLinear Change 

of Refractory Plastic and Ramming Mix Specimens 

(2) ASTM C832 (2005), Standard Test Method of Measuring the Thermal 

Expansion and Creep of Refractories Under Load 

(3) ASTM C1148 (2002), Standard Test Method for Measuring the Drying 

Shrinkage of Masonry Mortar (Not specifically for refractory concretemodified) 

(4) ASTM E228, Standard Test Method for Linear Thermal Expansion of 

Solid Materials With a Push-Rod Dilatometer 

(5) ASTM C 1171, Standard Test Method for Quantitatively Measuring the 

Effect of Thermal Shock and Thermal Cycling on Refractories 

e. Modulus ofRupture 

(1) ASTM C 133 (Reapproved 2008), Standard Test Methods for Cold 

Crushing Strength and Modulus of Rupture ofRefractories 

f. Thermal Conductivity 

(1) ASTM C 1113/C 1113M, Standard Test Method for Thermal Conductivity 

of Refractories by Hot Wire (Platinum Resistance Thermometer 

Technique) 

g. Other Physical Properties 

(1) ASTM.C 20 (Reapproved 2010), Standard Test Methods for Apparent 

Porosity; Water Absorption, Apparent Specific Gravity, and Bulk Density 

of Burned Refract~ry Brick and Shapes by Boiling Water 

(2) ASTM C 830 (Reapproved 2006), Standard Test Methods for Apparent 

Porosity, Liquid Absorption, Apparent Specific Gravity, and Bulk Density 

of Refractory Shapes by Vacuum Pressure 

(3) ASTM C 1419 (Reapproved 2009), Standard Test Method for Sonic 

Velocity in Refractory Materials at Room Temperature and Its Use in 

Obtaining an Approximate Young's Modulus 

18 
NASA/TM-2010-216293R 

(4) ASTM E 1269, Standard Test Method for Determining Specific Heat 

Capacity by Differential Scanning Calorimetry 

3.1 Sample Preparation 

Prior to use or test, refractory materials shall be stored in dry, protected, and weatherproof 

structures. In general, appropriate care should be exercised to protect the raw materials from 

weather and ensure that the manufacturer's recommendations are followed during mixing and 

installation. Raw material expiration dates should be checked prior to use to ensure that the final 

material has the properties advertised by the manufacturer and meets the requirements ofNASA. 

Care should be exercised to ensure that the correct product is used for a particular application. 

Refractory products are often formulated to cast or gun the refractory product. Refractory 

materials for laboratory testing should be prepared with the same method that would be used to 

install the material at the pad. 

In general, concrete samples can be prepared in three different ways: shotcreting, gunning, or 

casting. Casting involves mixing the components and pouring the cementitious product in place 

to cure. Shotcreting is an automatic delivery method that pumps a wet (already mixed) concrete 

to the nozzle where air is added to deliver the concrete to the target. Gunning is a process in 

which the cement and sand are injected into an air stream to deliver it to the nozzle. At the 

nozzle, the dry mix and water are combined, and the materials are pneumatically expelled to the 

target. 

3.1.1 Sample Preparation - Shotcreting or Gunning 

Forms for the production of refractory samples by the gunning process can be fabricated using 

ASTM C1140-03 (Appendix E). Bulk samples utilizing this specification have been prepared at 

KSC. See Endnote 7 in Section 8. In accordance with the requirements of ASTM C1140, 

samples cannot be taken from the bulk sample in the space equal to the depth plus 1 inch from 

the outside edges. The form should be constructed from rigid materials, so that dislodging of the 

refractory product through vibration or deformation is prevented. The walls and bottom of the 

forms fabricated at KSC used a 0.5-inch wire fabric mesh to contain the product. See Figure 10. 

19 
NASA/TM-2010-216293R 

2"x4" FRAME 

24"x24" WIRE 

FABRIC FORM 

(1/2" MESH) 

TOP VIEW 

BOTTOM 

r------ 24.0 ----; 

CROSS SECTION VIEW , 

Figure 10. Form~ Designed To Hold the Wet Refractory Concrete (Units in Inches) 

. . . 

To prepare samples, refractory materials shall be gunned or shotcreted into forms at an angle as 

shown i~ Figure 11. In the prior referenced testing, two forms were used for each sample in the 

study. Sample size or quantity should be sufficient to produce the amount of replicate samples 

needed for laboratory testing. After material placement, the samples shall cure under ambient 

conditions in accordance with 'manufacturers' recommendations. ASTM C 1140 recommends the 

samples be tightly wrapped to prevent water from evaporating during the curing process. 

-20 
----------

NASA/TM-2010-216293R 

HOruzONTALSURFACE 

Figure 11. Position of the Forms During Spray Application 

The gunning process and the formation of the bulk cementitious material have been reported (see 

Endnote 8 in Section 8) and photos of the sample preparation are shown in Figure 12. The photo 

on the left shows spraying of the refractory mix and the photo on the right shows gunned but still 

wet refractory material in the form. 

Figure 12. Sample Preparation (Zampell Refractories in Tampa, Florida) 

In accordance with ASTM C1140, the edges ofthe forms should be discarded, and the remainder 

of the sample should be cut to meet ASTM standards for sample testing. Each sample (cores and 

cubes) should be cut from the center interior section, discarding the top and bottom from each 

p1ece. 

21 
NASA/TM-2010-216293R 

To prepare individual samples for testing, the ambient-cured bulk refractory material is removed 

from the form and cut using both a diamond saw and a core drill. Figure 13 shows an example 

cut pattern that has been used to produce cylindrical, cube, and bar shaped samples that were 

required for laboratory testing as a part of a prior project. See Endnote 9 in Section 8. The dark 

areas on the side are discarded. 

r--------24"--------t 

24" 

Figure 13. 

. SCRAPAREA 

MORC133 

MORCI33 

Cll71 Cll71 

SCRAP AREA SCRAP AREA 

MOR Cl33- 5 SAMPLES (2"x2"x9", from center) 

C1171- 10. SAMPLES (1 "xl "x6", from center) 

Compression Cores- 20 SAMPLES (2" round x 2" tall, from center) 

Example of Cut Pattern for Each Sprayed Sample (Units in Inches) 

3.1.2 Sample Preparation - Casting 

Cast specimens should be produced following the guidance of ASTM C862. According to this 

procedure, a paddle mixer (Figure 14) is used to prepare the wet cementitious mixture. ASTM 

C862,gives details on water addition and mixirig procedures. 

As quickly as possible following the mixing phase, the wet cementitious material should be 

packed into appropriate molds for the tests to be conducted. The product is then consolidated 

. using a vibration table until the top surfaces appear smooth. The filled molds are then placed in a 

humidity chamber (greater than 95% humidity) at ambient temperature for a curing period of 24 

hours. The specimens are then removed from the molds and are allowed to continue curing under 

ambient conditions for at least 7 days. 

22 
NASA/TM-2010-216293R 

Figure 14. Paddle Mixer Used at KSC To Prepare Refractory Samples 

3.2 Laboratory Testing and Standards 

This section describes the tests that can be used to evaluate the performance of a refractory 

material in a controlled laboratory environment. These tests are suggested to be part of the new 

qualifying process of a refractory material, though they cannot be used to qualify a refractory 

product without physical launch pad (or simulated launch pad) testing. The official qualification 

procedure is found in Section 2. 

3.2.1 Sieve Analysis 

Particle size distribution of a refractory concrete can be explained using ASTM C92 (Note 13). 

This test method utilizes a set of calibrated sieves for the analysis prewetted and wetted 

cementitious material. According to this standard, the concrete is weighed prior to analysis and 

placed through a series of sieves, starting with the one with the coarsest opening. Results from 

this standard are based (and reported) upon the material retained in each sieve, as well as the dust 

loss of the material passing through the final sieve. 

23 
NASA/TM-2010-216293R 

3.2.2 Specific Heat Analysis 

Specific heat capacity, or simply specific heat, can be defined as the quantity of heat required to 

raise the temperature of 1 gram of substance by 1 degree Celsius at constant pressure. ASTM 

E1269 is used to determine specific heat using differential scanning calorimetry (DSC). DSC 

measures the specific heat of materials by measuring differences in the heat flow into the test 

material and a reference material or blank. 

The specific heat of the sample (as a function of temperature) is a reportable quantity that may be 

used by engineers to help select refractory products to satisfy launch pad requirements. 

3.2.3 Cold Compressive Strength 

The compressive strength of a refractory material can be determined in accordance with ASTM 

Cl33. In brief, the cold crushing strength or compression strength (CS in lbtfin2 or psia) is 

defined by: 

where, 

Fmax 

A 

applied (in2

) 

CS= 

Fmax 

A 

= maximum force or load applied at the yield point of the material (lbr) 

= average (between top and· bottom) cross-sectional area tq which the load is 

A complete description of the process used to measure cold crushing strength is in ASTM C133. 

Samples are usually 2-inch cubes or 2-inch-diameter cylinders. Prior to testing, the samples 

should be dried in a 110°C oven for 18 hours. The compressive strength can then be measured 

using standard mechanical or hydraulic compression testing machines conforming to the 

requirements of ASTM E4, Practices for Force Verification of Testing Machines. A photo of the 

Instron Universal Test Machine (Model5889) that meet these requirements is shown in Figure 

15. Figure 16 shows samples before and after crushing. · 

24 
NASA/TM -201 0-216293R 

Figure 15. Instron in the Material Testing Laboratories, O&C Building, KSC 

When tested according to ASTM C133 and according to the requirements ofKSC-SPEC-P-0012, 

the refractory material shall develop a minimum compressive strength of 31.0 megapascal (MPa) 

(4500 psi) at 7 days, with 90 percent of the 7-day strength developed within 24 hours. 

Figure 16. Cylindrical Sample Prior To and After Crushing 

3.2.4 Abrasion Resistance 

Solid rocket motors (SRMs) produce aluminum oxide (Ah03) and hydrochloric acid as a 

by-product. When the SRMs are operational, the particulates are expelled from the motors at 

considerable velocities. As a consequence, it is desirable to investigate the refractory materials 

25 
NASA/TM-2010-216293R 

resistance to abrasion. ASTM C704 is used to investigate this phenomenon under ambient 

laboratory conditions. 

Samples used for ASTM C704 are cut from bulk refractory materials and should measure 

between (4" by 4" by 1 ")to (412" by 412'' by 212'' or 3"). Prior to testing, the samples are dried in 

an oven at 110°C until a constant weight is achieved. The sample's resistance to abrasion is then 

investigated by propelling 1 OOOg silicon carbide media in air at a pressure of 65 psi. The abrasive 

media is impinged upon the sample at a 90-degree incident angle. 

An important point to note is that ASTM C704 investigates the abrasion resistance of refractory 

materials under ambient conditions. The abrasion resistance under elevated temperature launch 

pad conditions may differ greatly. 

3.2.5 Cold Modulus of Rupture 

The modulus of rupture can be explained in accordance with ASTM C133. The cold strength of a 

refractory material gives an indication of its suitability for the materials use in refractory 

construction, but it should not be construed as providing an equivalent level of performance at 

elevated temperature. The modulus of rupture (MORin lbtlin2 

or psia) is defined by 

where, 

3FmaxL 

MOR = 2bdA2 

F max = maximum force applied at rupture (lbr) 

L = span between supports (inches) 

b = breadth or width of the specimen (inches) 

d =:;; dypth of the specimen (inches) 

Typically, the modulus of rupture is determined from cast or gunned refractory material with 

nominal dimensions of9'; by 2" by 2". Prior to testing, all specimens should be placed into a 110 

oc oven until constant mass is achieved. Testing must be completed within 2 hours after the 

specimen is removed from the oven. 

The measurements are performed (with rectangular bars and three stress points) using a standard 

mechanical or hydraulic compression testing machine conforming to the requirements of ASTM 

E4. Examples of refractory specimens and load-bearing cylinders for MOR testing are shown in 

Figure 17 and Figure 18. Figure 17 shows the condition of a refractory sample prior to MOR 

testing. Figure 18 shows the same sample after the MOR process is complete. 

26 
--------------------- ---

NASA/TM -201 0-216293R 

Figure 17. Modulus of Rupture Testing- Fondu Fyre (Pad Sprayed), Sample 81, 

Control, MOR = 837 psia (Before Compression) 

Figure 18. Modulus of Rupture Testing- Fondu Fyre (Pad Sprayed), Sample 81, 

Control, MOR = 837 psia (After Compression) 

27 
NASA/TM-2010-216293R 

3.2.6 Thermal Expansion and Creep of Refractories 

The shrinkage\thermal expansion shall be tested in accordance with ASTM C832. In summary, 

this test method subjects refractory materials to elevated temperatures under a 25-psi 

compressive stress for 50 hours. During the process, sensors continuously measure the. linear 

change of the specimens parallel to the direction of the compressive stress. 

A review of prior sensor data indicates that the maximum temperature achieved on the launch 

pads is 2165 op atthe time of this writing. It is suggested that the thermal expansion and creep of 

the refractory product be measured to a temperature equivalent to the maximum temperature that 

the refractory concrete is subject to at the launch pad. Sensor data for the final Shuttle flights 

may necessitate that the refractory product is evaluated at higher temperatures. The thermal 

expansion under load shall be reported. The reader is encouraged to refer to the latest revision of 

KSC-SPEC-P-0012 for thermal expansion requirements. 

3.2. 7 Thermal Shock 

The thermal shock properties of a refractory material can be explained in accordance with 

ASTM C1171, ASTM C1419, and ASTM C133. These test methods are used to indicate the 

extent to which a refractory material can withstand stresses generated by sudden changes in 

temperature. 

ASTM C1171 relies on sudden changes in temperature to generate stress within the refractory 

materials. Crack and flaw-free samples with the nomimll dimensions of 1" x 1" x 6" are cut from 

the bulk material and are subsequently dried to a constant weight in a 110 °C oven. The sonic 

velocity is determined according to ASTM C1419, and the samples are then divided into two 

separate groups. Modulus of rupture testing (ASTM C133) on the first group (prior to heating) is 

used to make comparison to the post cyclic sonic velocity and modulus of rupture evaluations 

after the heating cycles. 

One heating cycle results from the exposure of the sample to an elevated temperature of 1200 ± 

· 15°C (2190 ± 25°F) for a 1 0-to-15 minute duration. The samples are then removed from the 

furnace for 10 to 15 minutes. This procedure is considered a single cycle. According to ASTM 

C 1171, this procedure (or cycle )"is repeated five times, the sonic velocity is again determined, 

and MOR testing is performed to evaluate thepost cyclic strength of the sample.· 

A direct comparison of velocity and strength loss measurements is then used to help delineate the 

performance of different refractory materials. Furthermore, any spalling of the specimens during 

the cyclic heating and cooling process should be noted and can be used as a metric to disqualify 

the use of the product. 

3.2.8 Thermal Conductivity 

The thermal conductivity of the refractory samples may be determined in accordance with 

ASTM C1113. In summary, ASTM C1113 uses a constant electric current that is applied to a 

pure platinum wire between two bricks of the refractory material. The heat that is generated from 

the electric current is conducted away fr.om the wire at a rate that is dependent upon the thermal 

28 
NASA/TM-2010-216293R 

conductivity of the refractory material. Using a minimum of four test temperatures, the k-Value 

(thermal conductivity) can be ascertained from this standard. 

3.2.9 Acid Resistance Tests 

Because of the presence of some amount of acid in the launch environment, or the potential 

thereof, the acid resistance,should be determined/compared for different candidate flame trench 

refractory materials. 

The acid soak test is used .to determine a refractory material's resistance to degradation when 

placed in a bath of0.1M hydrochloric acid (HCl). This procedure has been performed in the 

NASA Corrosion Technology Laboratory at the Kennedy Space Center. 

The test compares the cold compression strength before and after acid exposure to evaluate acid 

resistance. The test is performed as follows. Specimens consisting of 2-inch cubes or 2-inch tall, 

2-inch-diameter cylinders are dried overnight in a 110 °C oven. Each sample is individually 

submerged in a 200-ml volume ofthe 0.1M HCl acid in a plastic cup for 14 days. Midway 

through the exposure, the HCl is drained and a new solution is added. After 14 days, the 

specimens are rinsed and dried in an oven, and the compression strength is measured according 

to ASTM C133. The percent loss of strength after acid exposure is used to compare the relative 

acid resistance of different materials. 

3.2.10 Density; Porosity, Water Absorption, and Apparent Specific Gravity 

Porosity is the ratio of the volume of the open pores to the bulk volume of a material. Bulk 

density is defined as the mass of a material divided by the total volume it occupies. The total 

volume includes particle volume, interparticle void volume and internal pore volume. As with 

crushing strength, both water content and heat treatments factor significantly into apparent 

porosity and bulk density of the refractory material. See Endnote 10 in Section 8. 

Apparent porosity, water absorption, apparent specific gravity, and bulk density are primary 

properties of refractory materials that can be measured using ASTM C20 or ASTM C830. These 

properties are widely used in the evaluation and comparison of product quality and as part of the 

criteria for selection and use of refractory products in a variety of industrial applications. 

According to ASTM C20, refractory samples are dried prior to testing in a 110 °C oven until a 

constant mass is achieved. This mass is recorded. Specimens are then submerged in boiling water 

for a 2-hour duration and are allowed to cool. The weight of each sample is then measured while 

the samples are suspended in the water (to calculate the specific gravity). Finally, the samples are 

blotted dry with a moistened linen to remove all water from the surface, and then they are 

subsequently weighed. 

29 
NASA/TM-2010-216293R 

The bulk density of the samples is calculated as follows 

Where, 

B= bulk density 

D = dry weight 

V=volume 

D 

B =-

v 

The porosity of the samples is calculated as follows 

Where, 

P =porosity 

W = saturated weight 

D = dry weight 

V=volume 

W-D 

P,% = V X 100 

The water absorption of the samples is calculated as follows 

Where, 

A = water absorption 

W = saturated weight 

D = dry weight 

W-D 

A,%= D X 100 

The apparent specific gravity is calculated as follows 

Where, 

T =apparent specific gravity 

D = dry weight 

S = suspended weight 

D 

T = D -S 

30 
NASA/TM-2010-216293R 

3:2.11 Petrographic Analysis of Refractory Materials 

Concrete petrography is the study of hardened concrete microstructure using microscopic 

techniques. The study of concrete microstructure can be used to investigate properties and issues 

related to mix design, water/cement ratios, composition, degradation, chemical attack, corrosion 

of steel reinforcement, microcracking, porosity, and grain size. To investigate these issues, crosssectioned samples of (cured) refractory materials can be analyzed prior to and immediately after 

being exposed to rocket blast. 

In addition to optical microscopy, scanning electron microscopy (SEM) with energy dispersive 

x-ray spectroscopy (EDS) can provide information that is beneficial to the analysis of the 

cementitious materials. SEM/EDS is particularly useful in determining mineralogical 

composition and in visually documenting significant pores, pore throats, clays, framework 

grains, and cements. The NASA Corrosion Technology Laboratory and the NASA Materials 

Sciences Laboratory at the Kennedy Space Center have the equipment necessary to perform a 

suitable analysis. 

X-ray diffraction is an established technique that provides semi-quantitative determination of 

sample mineralogy and can be used to estimate and elucidate hydrated and nonhydrated phases 

of refractory materials. The NASA Corrosion Technology Laboratory and the NASA Materials 

Sciences Laboratory at the Kennedy Space Center have the equipment necessary to perform a 

suitable analysis. 

4 ROCKET ENGINE EXHAUST TESTING 

4.1 LC-39 Full-Scale Testing 

To meet the requirements of the standards in Section 2, a new refractory concrete product must 

be exposed to an actual launch environment. To assist in this requirement, two test stands were 

constructed and installed at the bottom east and west sides of the main flame deflector, SRB side 

(Figure 19). The east and west test stands are capable ofholding three samples each (Figure 20). 

31 
NASNTM-2010-216293R 

Figure 19. Typical Three Sample Test Stand 

Figure 20. Test Stand Locations 

Samples for this testing must be prepared in the same manner as they are intended to be used. 

The finished dimension of the test samples must be in accordance with sketch in Figure 21 , taken 

from KSC-SPEC-P-0012. 

32 
Figure 21. 

314' 

114"R 

TVPICAL SECTION 

4" 

NASA/TM-2010-216293R 

NOTE: 

SURF ACE EICPOSLO TO 

ROCKET ENGINl [)CHAUST 

REFRACTORY 

CONCMfTE 

llr X 14 GAUGE. 

GAIDSU£L 

REINFORCING 

All DIMENSIONAl TOlERANCES 

SHAll BE: tO 

-Ill INCH 

Typical Sample Size From KSC-SPEC-P-0012 

33 
NASA/TM-2010-216293R 

4.1.1 Installation and Prelaunch Analysis 

. . Based on the current effort to qualify additional refractory materials for the use on the LC-39 

main flame deflector, test samples shall be installed at designated areas of the facility and then 

subjected to rocket engine exhaust. The intended procedure is outlined below and is based on the 

historical testing completed at KSC. Specific instances of work will be coordinated by NE-M9 

and data transferred to NE-L for documentation by a KSC Materials Sciences Laboratory (MSL) 

report. 

The specimen shall conform to the dimensions outlined in the latest revision ofKSC-SPEC-P0012, and installed in a manner similar to the test frames required in Ground Support Equipment 

Engineering Order E03-79K09546, Refractory Test Frames. · 

Panels shall be installed with GE SCM 3404 on the bottom of the refractory concrete test panels 

to ensure levelness and adequate structural suppolj: of the integrated test panel. 

Tap test stall be performed by NE-M9 and United Space Alliance (USA) engineering to 

determine the soundness of the samples. 

Ablative coating, such as Dow Coming 3-6077 Silicone Ablative RTV (white colored) NSN 

10753-0012-801 or other approved equal per KSC-SPEC-F-0006, shall be applied to the areas 

surrounding the panels to protect the fixtures used to secure the samples. 

The installation shall be documented with photographs by NE-M9/NE-L or a designee. 

The dimensions of the test specimens shall be measured after installation to a tolerance of 1.5 

mm (1116 in) using an optical scanner with a higher resolution performed by USA Optics. 

4.1.2 Postlaunch Analysis 

. The postlaunch condition shall be documented with photographs by NE-M9/NE-L or a designee. 

A postlaunch survey of any liberated debris may be performed by base operations contractor 

based on a consensus within NE. 

The dimensions of the test specimens shall be measured after installation on the deflector to a 

tolerance of 1.5 mm (1116 in) using an optical scanner with a higher resolution performed by 

USA Optics. 

The test samples will not be removed unless the sample loses pieces of a 2 in x 2 in x 2 in or 

greater, or other structural detect is denoted by NE-M9/NE-L and USA engineering when 

determining the soundness of the samples. 

For multiple launches, the test stands will be swapped, with intact samples remaining installed 

between launches. 

34 
NASA/TM-2010-216293R 

The dimensions of the test specimens shall be measured after installation on the deflector to a 

tolerance of 1.5 mm (1/16 in) using an optical'scanner with a higher resolution performed by 

USA Optics. 

An additional sample coupon may be installed at the direction ofNE-M9 to avoid the possibility 

of generating FOD during launch. Sample installation will proceed in accordance with previous 

installation section. 

Installation shall be documented with photographs by NE-M9/NE-L or a designee. 

The sample may remain in test indefinitely at the direction ofNE-M9. 

4.2 Subscale Rocket Engine Exhaust Test Stands and Material Dimensions 

Subscale refractory testing has been performed at various locations. This section presents 

preliminary designs for refractory material sample size and test stand configurations using a 

standard 5-inch center perforated (CP) solid rocket motor. The basic designs are a modification 

of 80K60515 drawings used for full-scale live launches at LC-39A. The subscale samples are 

designed (smaller) for easier handling; 26.5-inch square at the base instead of29.5-inch square, 

resulting in a 20 percent reduction in volume. Figures 22 through Figure 26 depict the new test 

stand designs and material dimensions for subscale testing. 

35 
·NASA/TM-2010-216293R 

J 

1 

0 0 

6.0 

0 0 

6.0 

29.5 t 0 0 

6.0 

0 0 

6.0 

j 

0 0 ll 0 0 0 0 I 

f---6.0 6.0 - 6.0-

1-----------31.5--------------1 

Figure 22. Preliminary Test Stand Design, Top View 

36 

® 

_j 
DRILL & TAP FOR 

1/2"-13 BOLT 

I/2"-13x3" BOLT 

WI WASHER 

DRILL & TAP FOR 

1/2"-13 BOLT 

I/2"-13x2" BOLT 

WI WASHER 

NASA/TM-2010-216293R 

t-----------25.0 ----------t 

3/4" x 14 GA GRID 

STEEL REINFORCING 

1-------------31.5--------------1 

DETAIL 

SCALE:TBD 

1-----------25.75 -----------j 

3/4" x 14 GA GRID 

STEEL REINFORCING 

1------------29.5 ------------t 

DETAIL 

SCALE:TBD 

Figure 23. Cross-Section Details From Top View 

~r 24.0 

1/2" PLATE 

l/2"-13x3" BOLT 

W/WASHER 

>"CPT ROCKET MOTOR · J 

'-L---_--------L_______j STAND· SIDE VIEW 

Figure 24. Test Stand Side View 

37 
NASA/TM-2010-216293R 

26.5 ~..., I 

1---------25.0-----,+----11 

26.5 25.8 

L 

J REFRACTORY CONCRETE TEST SAMPLE 

SCALE:TBD 

Figure 25. Refractory 

L~ 25.0 f" 

"" 

"·" 

~3/4" x 14 GA GRID 

STEEL REINFORCING 

26.5 

DETAIL 

SCALE: TBD CB 

ri 25.8 f" 

" 

. "·" 

"""=3/4" X 14 GA GRID 

STEEL REINFORCING 26.5 

DETAIL 

SCALE: TBD ® 

Figure 26. Refractory Material Sample Cross-Sections 

38 
NASA/TM-2010-216293R 

5 SUBSCALE PARAMETERS 

Qualification testing for new refractory materials must take place in either an actual rocket 

launch or a rocket engine test (KSC-STD-164). This section provides environmental data taken 

during the launches ofSTS-126, STS-119, STS-125, and STS-127, which can be used to provide 

guidance on the test parameters of a particular subscale test. Temperature, pressure, acoustic 

pressure, acceleration, strain, total heat flux (calorimeter), and radiative heat flux (radiometer) 

were measured during those launches. In addition, environmental data found in KSC-GP-1059, 

KSC-DD-818-TR, and KSC-DM-3649 should be consulted when finalizing subscale test 

parameters. Temperature, pressure, and heat flux values are thought to be directly comparable 

between the Shuttle launch and sub scale testing. Maximum values of these parameters, as well as 

graphs that show the parameters over time, are reported. Strain and acceleration are impacted by 

the overall structural design and cannot be simulated in subscale testing. There are limitations to 

subscale testing, which will be discussed as well. 

5.1 Launch Environment Assessment 

The launch environment was measured in the flame trench during the launches ofSTS-126 

(Endeavour, November 14, 2008), STS-119 (Discovery, March 15, 2009), STS-125 (Atlantis, 

May 11, 2009), and STS-127 (Endeavour, July 15, 2009). Complete results are reported in 

NASA-TM-2010-216294, KSC Launch Pad Flame Trench Environment Assessment. 

Temperature, pressure, acoustic pressure, acceleration, total heat flux (calorimeter), and radiative 

heat flux (radiometer) were measured. The sensors used for each measurement and their 

locations are given in Table 4. The locations of the sensors are shown in Figure 27 thru Figure 

30. Sensors were located on both the east and west flame trench walls for locations 1 - 7. The 

specific location was labeled with an "E" or "W" after the number to denote the east or west wall 

as shown in the figures. Sensors at locations 1 - 4 were exposed to the exhaust environment. 

Locations 4E and 4W were at the bottom of the deflector near the walls and adjacent to the test 

stands and will be used to test materials during the last Shuttle launches, as shown in Figure 31. 

The sensors for accelerometer measurements at locations 5 - 7 were placed in the catacombs on 

the cold face of the flame trench wall. These sensors were not directly exposed to the heat of the 

launch but were used to analyze the acoustical response of the reinforced concrete structure. 

Locations 8W and 9W were located on the west side flame deflector, underneath the MLP 

(Figure 29). These locations did not have equivalent measurements on the east side. 

Accelerometer and strain gauge measurements were made on the SRB and SSME flame 

deflectors, with locations as shown in Figure 30. These locations are the only ones where 

measurements were taken on the deflector structure at the time of this writing. 

There are plans to have sensors mounted directly on the hqt face of the SRB side of the main 

flame deflector during the launch of STS-133 and STS-134. The current plan is to install a 

combination of similar sensors, witness rods, and specially designed slug-type calorimeters. 

These measurements should be consulted when designing a subscale test. 

39 
NASA/TM-2010-216293R 

Table 4. Sensors Used During Launch Environment Assessment 

Measurement Sensor Model Name/No. Locations 

Temperature NANMAC 9300 Erodible Thermocouple 1, 2, 3, 4 

Pressure Stellar Technology ST 150 1, 2, 3, 4, 8W, 9W 

Acoustic Pressure Kistler 6013C 4 

Accelerometer Wilcoxon 797L 5, 6, 7, SFUB, SSME 

Calorimeter Medtherm 64-2000-600-19-20054AT 1, 2, 3, 4 

Radiometer Medtherm 4TP-2000-600-23-200264AT 1, 4 

Strain CEA-06-125UR-350 SFUB, SSME 

MLP 

•2E 

E •6E 

• . ' .. . .. . .. . . ' . ~ . . . . . . . . 

Figure 27. Sensor Locations on the East Wall of the SRB Flame Trench 

40 
Figure 28. 

NASA/TM-20 1 0-216293R 

LP 

7W 

• 

Sensor Locations on the West Wall of the SRB Flame Trench 

Right SRB 

ExlwuatHole 

North 

t 

Bottom View of MLP 

(Looking Under MLP) 

,..,., •• ~tJOO 

--

Left SRB 

Exhaust Hole 

Orbiter 

Exhaust Hole 

Figure 29. Location of Pressure Sensor 9W Underneath the MLP 

41 
NASNTM-2010-216293R 

Typical Panel 

(VIeWed from lnler Flame 

Deflector) 

KFtXli\ 001A RSRB Flame IJBflecD Vllr.ltion 

I<FDG*. 001A RSRB Rne llelecD 5tr.lin (1) 

1<FDG*. 002A RSRB Rne llelecD 5tr.lin (2) 

KFDGA.003ARSRB F&.ne Delledllr Strain (3) 

SSM: l)r:llerV 

KFtXli\ 002A SSY: Rne Delecb" VlnliDn 

I<FDG*. 004A. SSME Rne llelecb Sbain (1) 

I<FDG*. 005A. SSME Flame llelecb Sbain (2) 

KFDGA. 006ASSME F&.ne Delledllr Strain (3) 

Figure 30. Sensor Locations on the SRB and SSME Flame Deflectors 

42 
NASNTM-2010-216293R 

Figure 31. Location 4E Is Highlighted in the Circle 

5.2 Parameters for Subscale Testing 

This section lists the environmental parameters measured during four Shuttle launches and may 

be used as guidance for designing subscale rocket exhaust tests. Temperature, heat flux, and 

pressure are thought to be parameters that can be closely mimicked in subscale rocket exhaust 

testing. Acceleration and strain are dependent on the configuration of the overall deflector 

system and would be difficult to duplicate in subscale testing. At the time this report was written, 

the harshest environment for temperature and heat flux was found at locations 4E and 4W. 

Location 8W had the highest measured pressure. Results from future measurements on the hot 

43 
NASNTM-2010-216293R 

face of the flame deflector should exceed these ,values (for testing) and should be used when 

designing a subscale test. The maximum measured temperature, pressure, and heat flux (as of 

this date) are given in Table 5. · · 

Table 5. Maximum Temperature, Pressure, and Heat Flux Measured During the 

Launch of STS-126, STS-119, STS-125, and ·sTS~l27 

Parameter Maximum value Location 

Temperature 2165 °F .· 4E. 

Pressure 99A psig ., 8W 

Total heat flux (calorimeter) · . 2493 btu/ft2

·sec 4W 

Radiant heat flu?'- (radiometer) 72 .. 1 btu/ft2

·sec 4W 

The maximum values for temperature and radiant heat flux were· relatively consistent for the 

launches, while the maximum pressure and total heat flux were highly variable. Currently, 

KSC-SPEC-P-0012 necessitates that the refractory material withstand a heat flux of3300 

btu/ft2·sec for approximately 10 seconds. This is the only environmental condition specifically 

called out in the specification. The value is consistent with this measUred value, but it will likely 

be low when compared to the heat flux measured on the deflector face. The time of exposure, 10 

seconds, is about twice as long as seen during a launch of the Shuttle. Further details are given in 

the following sections, while complete results are reported in the NASA/TM-2010-216294. 

5.2.1 Temperature · 

The maximum temperatures measured at locations 4E and 4W are presented in Table 6. The 

temperatures were higher on the ~ast wall than on the west wall. Location 4 is the hottest 

measured point on both walls, reaching m~ximum temperatures of 2165 and 1922 op on the east 

and west sides, respectively. 

Table 6 .. Maximum Temperatures (°F) Measured During Three Launches 

Sensor Location STS-119 STS-125 STS-127 All Launches Max 

4E 1936 °F 2165 °F 21(}5 °F '\ 

4W 1654 °F 1806 °F 1922 °F 1922 .°F 

44 
NASA/TM-2010-216293R 

Figure 32 shows the temperatures at locations 4E and 4W during the launch ofSTS-125. At both 

locations, the temperature exhibits a double maximum. The two maxima occur about 2 and 5 

seconds after the initial heating. This behavior is consistent with the measurements at 4E and 4W 

during all launches. At the other locations, a single temperature maximum is reached at about the 

same time as the second maximum at location 4. This behavior was consistent during the 

different launches. 

2500.00 .-------------------------

--STS-125 Location 4E 

--STS-1 ~ Location 4W 

0 2 4 6 8 10 

Thne(sec) 

Figure 32. Selected Temperature Measurements During the Launch of STS-125 

5.2.2 Pressure 

Table 7 gives the maximum and minimum pressures at locations 8W, 4E, and 4W for each 

launch. The maxima and minima over all launches for each location are also given. As would be 

expected, the highest pressures were found on the side flame deflector, location 8W. At locations 

4 and 8, the pressure initially spikes for 0.2 to 0.4 second. After this, the reading is elevated, and 

it is consistent with the profile from other launches. At locations 1, 2, and 3, there is an initial 

increase in pressure, followed by a period of negative pressure, before returning to ambient 

atmospheric pressure. Figure 33 shows pressure data taken during STS-126 for location 8W, 4W, 

and4E. 

45 
NASA/TM-20 1 0-216293R 

Table 7. Maximum and Minimum Pressures (psig) Measured During Three Launches 

STS-126 STS-119 STS-125 All Launches 

Sensor location Max Min Max Min Max Min Max Min 

8W 91.5 -2.3 99.4 -2.1 56.4 -2.1 99.4 -2.3 

4E 45.6 -2.9 36.8 -22.9 61.1 -0.6 61.1 -22.9 

4W 45.1 -2.4 45.6 -3.1 42.1 -0.1 45.6 -3.1 

100 

- STS-126 Location 8W 

80 - STS-126 Location 4E 

-

1:)1) 60 ·- - STS-126 Location 4W ~~------------- ----------------- 'I) 

...._, c. 

e 40 

= 'I) 

'I) 

~ 

20 ~ "" 

0 

-20 

0 2 4 Time (sec) 6 8 10 

Figure 33. Selected Pressure Measurements During STS-126 

5.2.3 Calorimeter and Radiometer 

Total heat flux and radiative heat flux were measured with a calorimeter and radiometer during 

two launches. Total heat flux was taken at locations 1 through 4 on both walls, while radiative 

flux was taken only at locations 1 and 4. The convective heat flux can be inferred by taking the 

difference between the two values. 

Table 8 and Table 9 lists the maximum total heat flux measurements for locations 1 and 4 on 

both walls. The total heat flux was highest at location 4 W. At this location, the flux has two 

broad maxima near 2 and 5 seconds after initial heating begins. This is the only location to 

exhibit this behavior. At all locations, the heat flux spikes to large values for short periods, 

ranging from 10 to 100 milliseconds, during the launch. Radiative heat flux is also highest at 

location 4W, and it has the double maxima similar to total heat flux. Radiative heat flux does not 

46 
NASA/TM-2010-216293R 

exhibit the spiking behavior that total heat flux does. Figure 34 shows selected calorimeter 

measurements. Figure 35 shows selected radiometer measurements. 

Table 8. Maximum Calorimeter Measurements (btu/ft2 sec) Obtained 

During Two Launches 

Sensor Location STS-126 STS-119 All Launches Max 

IE 95.0 109.7 109.7 

4E 306.8 155.0 306.8 

1W 104.1 107.1 107.1 

4W 652.2 2492.5 2492.5 

3000.00 

- STS-119 Location 4E 

2500.00 

-

CJ - STS-119 Location 4W ~ 

riJ 

N 2000.00 

¢:: 

........ 

= .... 

.c 1500.00 -

~ 

-~ 

.... 

~ 1000.00 ~ 

= 

500.00 

0.00 

0 2 4 6 8 10 

Time (sec) 

Figure 34. Selected Calorimeter Measurements During STS-119 

47 
NASA/TM-2010-216293R 

80.00 

70.00 

-

CJ 60.00 ~ 

- STS-126 Location 1E 

Ill 

.... 

¢:: 50.00 

= - STS-126 Location 4W 

.... 

,l:J 40.00 -~ 

= 

-~ 30.00 .... 

~ 

~ 

= 20.00 

10.00 

0.00 

0 2 4 6 8 10 

Time (sec) 

Figure 35. Selected Radiometer Measurements During STS-126 

5.3 Limitations of Subscale Testing 

The measured values of temperature, pressure, and heat flux reported here are not the highest 

values the refractory material will see, since the sensors were not in the most severe locations. 

Future launches will attempt to measure these parameters directly in the SRB impingement zone. 

Before conducting a subscale test, these updated values and sensor locations should be taken into 

account. 

Subscale testing is likely to cause more erosion than would occur during full-scale launch 

because of the concentrated rifling effect of the smaller rocket motor. Earlier subscale testing, 

reported in an AIAA 92-3978, ASRM Plume Deflector Analysis Program, found a smaller 

BATES aluminum oxide particle heat flux about three times larger than for the full-scale ASRM 

motor plume, because of the much smaller scaled radius. 

It is not possible to duplicate the launch environment seen at LC-39A and 39B. Strain and 

acceleration are dependent not only on the characteristics of the rocket motor exhaust but on the 

design of the launch structure as well. In addition, all parameters of a subscale test may not 

directly correlate to full-scale launches. 

Even if a material withstands a subscale launch test, it may not be successful in full-scale 

launches. The structural design of the flame deflector and trench, the reinforcement of the 

structural material, and other environmental conditions (e.g., weather) all affect refractory 

material performance. 

48 
NASA/TM-2010-216293R 

6 ENGINE TEST STAND FACILITIES 

There are a few rocket engine test stands with the potential to perform subscale testing. Facilities 

at John C. Stennis Space Center (SCC), Marshall Space Flight Center, and A TK Aerospace 

Systems are described here. Many factors go into the selection of a subscale test facility. These 

factors include the size of the rocket, the use of liquid or solid fuels, the ability to scale the test, 

and the cost of using the facility. 

6.1 John C. Stennis Space Center 

Stennis Space Center (SSC) is located in Hancock County, Mississippi, at the Mississippi/ 

Louisiana border. It is NASA's largest rocket engine test facility. It is the primary center for 

testing and flight certifying rocket propulsion systems for the Space Shuttle and future 

generations of space vehicles. 

Construction of the 13 ,500-acre complex began in October 1961. The test area is surrounded by 

· a 125,000-acre acoustical buffer zone. The facility's large concrete and metal test stands were 

originally used to test-fire the first and second stages of the Saturn V rockets and are now used to 

flight certify the Space Shuttle Main Engines. 

The site was originally selected by the U.S. Government because it was located in a thinly 

populated area that had barge access. Furthermore, the site is advantageously located between 

the Michoud Assembly Facility and the launch facility in Cape Canaveral Air Force Station in 

Florida. SCC maintains a number of rocket motor test facilities with various capabilities of 

testing. 

6.1.1 A-1 and A-2 Test Stands 

Both test stands are a single-position, vertical firing fixture that can accommodate test articles up 

to 33 feet in diameter. An exterior view of the A-2 test complex is shown in Figure 36. 

49 
NASA/TM-2010-216293R 

Figure 36. Exterior View of the A-2 Test Stand 

Both test facilities are designed to use liquid hydrogen (LH2) and liquid oxygen (LOX) 

propellants and can accommodate support fluids that include gaseous helium (GHe), gaseous 

hydrogen (GH2), and gaseous nitrogen (GN2). The A-2 Test Stand is equipped with an altitude 

diffuser used to simulate altitude conditions during engine testing. The maximum dynamic load 

that each structure is capable of testing is 1.1 million foot-pounds. See Endnote 11 in Section 8. 

Figure 37 shows the relative size of the flame deflector in relation to members of the refractory 

site review team. 

50 
NASA/TM-2010-216293R 

Figure 37. View of the A-2 Flame Duct and Diffuser 

6.1.2 B-1 and B-2 Test Stands 

Each B-1 and B-2 Test Complex consists of a dual-position, vertical, static-firing test stand. The 

B Complex is 295 feet tall and is equipped with a 200-ton main derrick-lifting crane. The test 

stand was designed to use LH2 and LOX propellants and can accommodate various support 

fluids that include gaseous helium, gaseous hydrogen and gaseous nitrogen. The maximum 

dynamic load that each structure is capable of testing is 11 million foot-pounds. See Endnote 12 

in Section 8. 

The refractory material team from KSC intended to inspect one of the test stands. Unfortunately, 

a direct inspection at that time was impossible since the area was cleared for a test. The rocket 

test is shown in Figure 38. 

51 
NASA/TM-2010-216293R 

Figure 38. Rocket Motor Test at Stennis B-1 Complex 

As a part of a prior project, NASA Corrosion Test Laboratory personnel photo-documented the 

B-2 Test Stand flame deflector (Figure 39). 

Similar to the A-2 Test Stand discussed in 6.1.1, the B-2 test stand is a steel structure that is 

much larger in scale. Figure 39 is used to give a comparative size ofthe B-2 flame duct in 

relation to others that are discussed in this report. 

52 
NASA/TM-20 1 0-216293R 

Figure 39. Stennis B-2 Flame Deflector 

6.1.3 E-1 Test Facility 

The E-1 Test Facility, originally designed as a developmental rocket engine component test 

facility for the National Launch System (NLS) Program, is available for developmental testing 

projects requiring high pressure and high flow rate cryogenic fluids, hydrogen, oxygen, inert 

gases, and industrial water. The E-1 facility is a multicell (3) with horizontal testing capabilities 

for large-scale propulsion programs (Figure 40). 

Figure 40. Stennis E-1 Flame Deflector 

53 
NASA/TM-2010-216293R 

The E-1 , Cell 1 is primarily designed for pressure-fed L02/LH2, L02/HC, and hybrid-based 

motors with thrust loads up to 750K pound-force. TheEl , Cell2 is designed for LH2 and L02 

turbopump assembly testing with thrust loads up to 60K pound-force. The E1 , Cell3 is designed 

for L02-rich turbopump assembly testing with thrust loads up to 60K pound-force. 

6.1.4 E-2 Test Facility 

The E-2 Test Facility was constructed to support materials development by subjecting test 

articles (including refractory concrete) to extreme temperature conditions and fluctuations. This 

facility has support capabilities, which include hot gas, cryogenic fluids, gas impingement, inert 

gases, industrial gases, specialized gases, hydraulics, and water. 

The E-2 facility is a multi cell complex that is capable of testing intermediate-size engines in both 

the vertical and horizontal configurations (Figure 41 ). 

Figure 41. E-2 Celll Test Stand at Stennis Space Center 

E-2 Cell 1 is a horizontal test cell, and utilizes propellants such as LOX, LH2 and RP-1. The 

horizontal test cell is capable of testing motors with thrust loads up to 120 thousand foot-pounds. 

See Endnote 13 in Section 8. Cell 2 is the vertical test cell and utilizes LOX and RP-1 

propellants. This test cell is capable of testing motors with up to 100 thousand foot-pounds of 

thrust. See Endnote 14 in Section 8. 

Figure 42 shows an overview of the Cell 2 flame duct. As shown in the photo, the flame 

deflector was built from steel over an I-beam.steel structure. The facility has a 4,000 gallon per 

minute (gpm) water deluge system that is used to protect the flame duct from plume radiant 

heating during testing. See Endnote 15 in Section 8. 

54 
------------

NASA/TM-20 1 0-2l6293R 

Figure 42. Stennis Space Center E-2 2 Vertical Test Flame Duct 

6.1.5 E-3 Test Facility 

The E-3 horizontal test cell can test motors up to 60 thousand foot-pounds of thrust and has 

support capabilities that include LOX, G02, and GH2. Cell2 is a vertical test cell that is capable 

of testing engines that use LOX, hydrogen peroxide, and HC propellants. Cell 2 can 

accommodate engines with thrust loads up to 25 thousand foot-pounds of thrust. See Endnote 16 

in Section 8. It can also be configured horizontally for using a Diagnostic Test Facility (DTF) 

rocket engine. This engine is used to generate a sub scale (approximately 5% of full scale) rocketdeflector test bed that has a similar plume impingement environment (heat flux, pressure, turning 

angle, etc.). Instrumentation measures the time-dependent erosion rates of various refractory 

materials and visualizes the plume/deflector interactions. 

Figure 43 shows the flame duct for the vertical E-3 test fixture. Numerous tests have been 

conducted at the vertical E-3 test stand. Tests have included small-scale combustion devices, 

such as catalyst beds, to larger devices such as ablative thrust chambers and a flight-type engine. 

See Endnote 17 in Section 8. 

55 
NASA/TM-2010-216293R 

Figure 43. E-3 Vertical Configuration Test Cell Flame Duct 

Figure 44 shows the horizontal E-3 test fixture. This configuration was set up specifically for 

testing erosion rates of candidate refractory materials. 

56 
NASA/TM-2010-216293R 

Figure 44. E-3 horizontal Test Configuration 

6.1.6 Building 3300 at Stennis Space Center 

Building 3300 at Stennis Space Center contained remnants of components used for scale model 

testing of candidate refractory materials. Examples of these components are shown in Figure 45 

and Figure 46. 

Figure 45. Test Duct at Stennis Space Center 

57 
NASNTM-2010-216293R 

Figure 46. Refractory Concrete Apron at Stennis Space Center 

These components were used as a part of a program designed to support the acquisition of data 

for baseline deflector design and refractory economical requirements. The program had four 

objectives: 

• Establish ROM bounds on the extent of material loss and damage. 

• Develop comparative data on the ability of various refractory materials to withstand the 

rocket plume environment. 

• Develop engineering and scientific data characterizing surface and plume interaction 

phenomena. 

• Evaluate (model scale) the operational capability of the deflector. 

Scaling of the test articles was driven by the availability of the Bates motor and propellant 

cartridges. The Bates motor (approximately 6000 lb thrust) was fired down the apron in a manner 

geometrically similar to that for a full-scale deflector. The plume deflector was designed in the 

configuration shown in Figure 4 7. See Endnote 18 in Section 8. 

58 
NASNTM-2010-216293R 

Figure 47. Stennis Plume Deflector Test Rig 

6.2 George C. Marshall Space Flight Center (MSFC) 

The MSFC is the U.S. Government civilian rocketry and spacecraft propulsion research center. 

The original home ofNASA, Marshall is today the agency's lead center for Space Shuttle 

propulsion and its external tank. Located on the Redstone Arsenal near Huntsville, Alabama, 

MSFC is named in honor of General of the Army George Marshall. Marshall's East Test Area 

contains various explosionproof test cells, separated from each other and equipped with basic 

propellants for testing full- and subscale rocket engines. The Advanced Engine Test Facility in 

Marshall's West Test Area is another area in the test laboratory used to test rocket engines. It is 

used to assess and validate new propulsion technologies and prototype hardware for large rocket 

engines. See Endnotes 18, 19, 20, 21, 22, and 23 in Section 8. 

6.2.1 Dynamic Test Stand 

The dynamic test stand is located in the East Test Area and is officially a Historic National 

Landmark. This test stand was built in 1964 for mechanical and vibration tests on fully 

assembled Saturn V rockets and was modified in 1977 for vibration tests on the mated Space 

Shuttle and for evaluation of the craft's dynamic characteristics. See Figure 48. 

59 
NASA/TM-2010-216293R 

Figure 48. MSFC Dynamic Test Stand 

6.2.2 Load Test Annex Facility 

Located in the East Test Area, this structural test stand features a multimillion pound, movable 

crosshead weight, mounted on four towers. The facility is capable of sustaining the force loads 

experienced by large launch vehicles and can accommodate stages of up to 100 feet high and 54 

feet in diameter. See Figures 49 and 50. 

60 
NASA/TM-2010-216293R 

Figure 49. MSFC Test Stand 116, Solid Fuel Torch Test 

Figure 50. MSFC Test Stand 115, Solid Fuel Torch Test 

6.2.3 Advanced Engine Test Facility 

The Advanced Engine Test Facility is located in the West Test Area and was originally designed 

for testing the first stage of the Saturn V moon rockets. The rockets boasted five F-1 engines, 

which produced a combined thrust of 7.5 million pounds. See Figure 51. 

61 
NASA/TM-2010-216293R 

Figure 51. A Scaled-Down 24-Inch Version of the Space Shuttle's Reusable Solid Rocket 

Motor Fired at a MSFC Test Stand 

6.3 ATK Aerospace Systems 

ATK is a pioneer in solid rocket propulsion systems, strategic missiles, missile defense, 

lightweight space deployables, solar arrays, and satellite thermal management systems. ATK's 

space propulsion and ordnance products reflect more than 45 years of experience, providing 

high-performance and reliable propulsion for the aerospace industry. Their capabilities are 

focused mainly on solid rocket motors. SRMs range in size from full-scale Space Shuttle 

reusable solid rocket motor (RSRM), 126 feet long and 12 feet in diameter, to smaller sizes such 

as 3-inch STAR series. Rocket motor test stand configurations are either in the horizontal or 

vertical position, depending on motor size, and can be modified according to test requirements. 

Small-scale testing can be performed on thermal protection materials as short as approximately 1 

second in duration. A typical 5-inch CP test would run approximately 2 to 3 seconds. Larger 

FPCs (40-pound charge) and SPCs (70- pound charge) motors configured horizontally will bum 

for up to approximately 35 seconds for the FPC and approximately 70 seconds for the SPC. Bum 

times and temperature depend on chosen propellant and throat diameter. See Figures 52, 53, and 

54. 

62 
-- - --------------- ----- - -------------- -------

NASA/TM-2010-216293R 

Figure 52. Horizontal Testing of an ASAS 21-120 Motor 

Figure 53. Space Shuttle RSRM 

63 
NASA/TM-2010-216293R 

Figure 54. 

7 LIST OF REFERENCES 

Space Shuttle RSRM Test Firing at the ATK, 

Promontory, Utah Facility 

AIAA 92-3978, ASRM Plume Deflector Analysis Program, Petersen, T.V.; Makel, D. B.; 

Thurman, C. 

ASCE 7 Guide, Guide to the Use of the Wind Load Provisions of ASCE 7-02?? 

ASTM C20, Standard Test Methods for Apparent Porosity, Water Absorption, Apparent Specific 

Gravity, and Bulk Density of Burned Refractory Brick and Shapes by Boiling Water 

ASTM C24, Standard Test Method for Pyrometric Cone Equivalent (PCE) of Fireclay and High 

Alumina Refractory Materials 

ASTM C92, Standard Test Methods for Sieve analysis and Water Content of Refractory 

Materials 

ASTM Cl33, Standard Test Methods for Cold Crushing Strength and Modulus of Rupture of 

Refractories 

ASTM Cl79, Standard Test Method for Drying and Firing Linear Change of Refractory Plastic 

and Ramming Mix Specimens 

ASTM C704, Standard Test Method for Abrasion Resistance of Refractory Materials at Room 

Temperature 

64 
NASA/TM-2010-216293R 

ASTM C830, Standard Test Methods for Apparent Porosity, Liquid Absorption, Apparent 

Specific Gravity, and Bulk Density of Refractory Shapes by Vacuum Pressure 

ASTM C832, Standard Test Method of Measuring the Thermal Expansion and Creep of 

Refractories Under Load 

ASTM C862, Standard Practice for Preparing Refractory Concrete Specimens by Casting 

ASTM C1113, Standard Test Method for Thermal Conductivity ofRefractories by Hot Wire 

(Platinum Resistance Thermometer Technique) 

ASTM C1140, Standard Practice for Preparing and Testing Specimens From Shotcrete Test 

Panels 

ASTM C1148, Standard Test Method for Measuring the Drying Shrinkage of Masonry Mortar 

ASTM C 1171, Standard Test Method for Quantitatively Measuring the Effect of Thermal Shock 

and Thermal Cycling on Refractories 

ASTM C1419, Standard Test Method for Sonic Velocity in Refractory Materials at Room 

Temperature and its Use in Obtaining an Approximate Young's Modulus 

ASTM E4, Standard Practices for Force Verification of Testing Machines 

ASTM E228, Standard Test Method for Linear Thermal Expansion of Solid Materials With a 

Push-Rod Dilatometer 

ASTM E1269, Standard Test Method for Determining Specific Heat Capacity by Differential 

Scanning Calorimetry 

KSC-DD-818-TR, Summary of Measurement of KSC Launch-Induced Environmental Effects 

(STS-1 Through STS-11) 

KSC-DE-512-SM, Facility, System, and Equipment General Design Requirements 

KSC-DM-3649, Lift-Off Response Spectra to Space Shuttle Launch-Induced Acoustic Pressures 

KSC-GP-1059, Environmental and Test Specification Levels Ground Support Equipment for 

Space Shuttle System Launch Complex 39 

KSC-SPEC-P-0012, Refra~tory Concrete, Specification for 

KSC-STD-164, Environmental Test Methods for Ground Support Equipment, Standard for 

KSC-STD-G-0003, Launch Support and Facility Components, Qualification of, Standard for 

KSC-STD-Z-0004, Structural Design, Standard for 

NASA-STD-5005, Standard for the Design and Fabrication of Ground Support Equipment 

65 
NASA/TM-2010-216293R 

NASA Technical Memorandum 83473, Terrestrial Environment (Climatic) Criteria Guidelines 

for Use in Aerospace Vehicle Development 

NASA-TM-2010-216294, KSC Launch Pad Flame Trench Environment Assessment 

NHB 7320.1, Facilities Engineering Handbook 

NSTS -7700, Volume XIV, Space Shuttle System Payload Accommodations 

SW-E-0002, Ground Support Equipment General Design Requirements 

8 ENDNOTES 

1. Coburn, S., "Atmospheric Corrosion," American Society for Metals, Metals Handbook, 

Properties and Selection, Carbon Steels, Metals Park, Ohio, 1978, 9th ed., Vol. 1, p. 720. 

2. Launch Complex 39A and 39B; http://science.ksc.nasa.gov/facilities/lc39a.html; Last 

accessed on December 17, 2008. 

3. Launch of Apollo 11; http://history.nasa.gov/ap11annlkippsphotos/KSC-69PC-442.jpg; Last 

accessed on September 28, 2010. 

4. http://science.ksc.nasa.gov/shuttle/technology/sts-newsreflstsover-prep.html#stsover-sound; 

Last accessed on March 8, 2010. 

5. http://spaceflight.nasa.gov/shuttle/reference/basics/launch.html; Last accessed on 

March 8, 2010. 

6. Darrow, E.A., Lays, E.; "Solid-Propellant Rocket Exhaust Effects (SPREE) and Methods of 

Attenuation - Volume 1: Project Summary," Report Martin-CR-65-93, Martin Marietta 

Corporation, January 1966. 

7. Perusich, S.; Sampson J.; Curran, J.; Hintze, P.; Whitten, M.; Parlier, C.; "New Refractory 

Materials for KSC Launch Pads 39A and 39B," NASA/KSC Internal Report (2009). 

8. Perusich, S.; Sampson J.; Curran, J.; Hintze, P.; Whitten, M.; Parlier, C.; "New Refractory 

Materials for KSC Launch Pads 39A and 39B," NASA/KSC Internal Report (2009). 

9. Perusich, S.; Sampson J.; Curran, J.; Hintze, P.; Whitten, M.; Parlier, C.; "New Refractory 

Materials for KSC Launch Pads 39A and 39B," (2009). 

10. Calle, L.M.; Kolody, M.; Curran, J.; Back, T.; Hintze, P.; Parlier, C.; Sampson, J.; Perusich, 

S.; Buchert, C.; "WBS 5.3.2 Commercial Offthe Shelf(COTS) Refractory Material 

Evaluation Report" (2009). 

11. A-2 Test Stand. http://sscfreedom.ssc.nasa.gov/esd/ESDTestFacilitiesA2.asp. Last accessed 

on September 1, 2010. 

66 
~~~~~~~

NASA/TM-2010-216293R 

12. B-1 Test Stand. http://sscfreedom.ssc.nasa.gov/esd/ESDTestFacilitiesBl.asp. Last Accessed 

September 1, 2010. 

13. Test Facilities Capability Handbook; Stennis Space Center; NP-2001-11-00021-SSC; 

November 2001. 

14. E-2 Test Facility. http://sscfreedom.ssc.nasa.gov/esd/ESDTestFacilitiesE2.asp. Last accessed 

November 7, 2008. 

15. Test Facilities Capability Handbook; Stennis Space Center; NP-2001-11-00021-SSC; 

November 2001. 

16. E-3 Test Facility. http://sscfreedom.ssc.nasa.gov/esd/ESDTestFacilitiesE3.asp. Last 

accessed November 7, 2008. 

17. Jacks, T.E.; Beisler, M. Expanding Hydrogen Peroxide Propulsion Test Capability at 

NASA's Stennis Space Center E-Complex; 39th AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit; 20-23 July 2003, Huntsville, Alabama. 

18. Douglas, F.; Dawson, M.C.; Orlin, P.A. ASRM Subscale Plume Deflector Testing; 

AIAA 1 ih Aerospace Ground Testing Conference; July 6-8, 1992; Nashville, TN. 

19. Sauvel, J. Static Test Defector for Ariane-6 SRM: Technical Design and Economic Choice. 

Proceedings of the second European Conference on Progress in Space Transportation. 

May 22-24, 1989. 

20. Orlin, P , F.; Dawson, D.M.; Bourgeois, S.; ASRM Plume Deflector Analysis; Sverdrup 

Technology, Inc.; Report No. 3112-92-016.; NASA/SSC; March 1992. 

21. Douglas, F.; Subscale Test Measurement Data Accuracy; Stennis Space Center Report 

No. 3112-92-013; NASA/SSC February 1992. 

22. ASRM Subscale Deflector Test Report. Vol. 1: Rough Order ofMagnitude (ROM) Testing. 

Sverdrup Technology, Inc. Report No.311292-008 NASA/SSC; December 1991. 

23. http://marshallstar.msfc.nasa.gov/7-20-06.pdf, Last accessed on September 23, 2010. 

67 
NASA/TM-2010-216293R 

This page intentionally left blank. 

68 
NASA/TM-2010-216293R 

APPENDIX A. MMA-1918-80, STS-1 

69 
NASA/TM-20i0-216293R 

! ~ 

/ 

·.1, 

--· ·-····-·-·-.. --T-A-SK-RE_Q_U_,_E_:~· '"& --1.-,-?~-,~-~-~M' J I -;i;·-- ···-------1 

I 

~ •va ·• · ' One ·Month after Launch of ~ t, D"T• •u•NITT•D r·· D·II"&D COM .. I.IICTIQN OATIC 

. ' ·-""Sept. 29 • 1980 : STS·l · · •·IAMIIL.K OCICIIIII~?IGNl . 

. .. 'Refractory Concrete Test Samples:.· ' . 1. Design Concrete Company ·. . , .. (a) Fondu.Fyre WA-1 (Approved·for:saturn) • 2 Test Samples 

(b) Fondu Fyre WA-l (Experimental) - 2 Test Samples 2. Wahl Refractory Products. Company · · (a) ·H. T~ Bond. Mortar·(Experimental)·· 3 Test Samples 

3. Harbfson-Wal ker Refractories . .... · (a) Refractory Concrete ----··-------· 2 Test Samples ~ · __ .__.,_ ) 

. (b) Refractory Concrete (Stainless Steel Fibers) - 2 Test Samp'l .. es ., ./ 

Samples for test - Refractory Concrete. 

Location to be SRB Flame· Deflector. ·79K09546. 

I• AHAL.YII. M•QuC•T&UI 

SEE 11REFRACTORY MATERIALS TEST PLAN" - attached, 

./ 

The three materials are on hand and testing is to be 1n1t1ated·by taking thickness measurement~·and color photographs of the initial condition of·the surfaces, 

) After exposure to SRB exhaust during STS-1 the samples are to be removed and · .. · returned to the laboratory for further thickness measurements and a· determination of 

surface erosion. The final surface condition such as spal11ng .. w111 be recorded and 

color photographs taken for comparison w1th initial surface conditions. 

An evaluation of performance and acceptance or ·rejection reconmendations will be 

made by formal report. 

• ... MONih II 1 MAI&. CODIII.I 

W, C1aut1ce~ PRC-1~7 867-3243 . PRC PRC-1217 

-t ............. ~~·.. p_g,~ 

'M .."/ffO " - ""- '··· . 6'! ·o· sen .. 

1 .. IIIHOHIIh ••• MAla,. CaUIIt Ia, UAT"I 

867·3748 · DD-MED·l 9/29/80 

FOR LAII IJU ONL'I 

·~!IN~·~··=·~=I~~~~;~:~t.n~~~~1~-~-··----~~~T-A~-K-~_-_;_·: __ /_9_1_1 __ ~_~_· _______ .1,'_ .._MJO_~-~-N-UM_._._~ __ ._. __________ ~ IUC I'DRM ll••t 'MKY, 1170, 

70 

I 

I 

.. 'l 

I 

~' J 
~ . 

. 

.. o· 

,;. 

R~FRACTOilY COl/CRETE ~IATERIALS TEST PLAII 

NASA/TM-2010-216293R 

·.·-·\ . I 

i 

I 

·I 

l. · MATERIALS FOR TESTIIIG :( KSC-SPI!C·P-0012) I 

I 

,.. ~ . 

1. Dosign Concretes Comp;my 

a. ·r-ondu Fyro WA-1* - 2 Test Sampl os · . b. Fondu Fyre WA-1 (Experimental)- 2 Test Samples 

Delivery date July 21. 1960 

.*Approved for Saturn 

'. 

2 • . ·wahl Refractory Products Company · ,· : . . a. II.T. Oond Hortar (experimental)- 3 Test Samples ' ' . ' 

I, 

1 

.I 

I 

I 

j 

I 

I 

The 3 test samples are at KSC;· · I 

The vendor delivered them on Feb. 12, 1960 . 1 

. They are stored by TG-FLD-.22 •. ·· . , ·1 

.' 3~ Harbison -Walker Refractories: ·: . . .. i I ; . • . • , • • · . • ·:1 

r

. -- ·-.- ------·a. -·Refractory Controto.- 2 Test. Samples- •...•.. -· -- -·-· -- -- .• ~ ··- :_: - -· -·~ · ·. [ b. ·Refractory Concrete - 2 Test Samples (Stai nl ass Stool Fif.lers) : . · 1 

[_ ·(\~a,-@ ::~:~~;:~:~J~;:~~:i;~~~-c~.~~- . . .. .. ----.. -· ... ·.. .~ . -. . . :. ···ll 

.a. Narco Cast 60.- 2 Test Samples ... 'b, .Narco Tab·- 2 Test Samples 

D_el1very date Aug. 1, 1980 . 

. ' I I.: ... TEST 'SAMPLE DELIVERY I HAtiDLIIIG AND STORAGC I 

The re.fractory concrete test samples a~ specified in KSC-SPEc;.p.oo12~: Figure .1. 

are. mass1ve,and requirc.'special handling, They .. we1'gh".approx1mately 200 lbs. · :each and are 2 Ft-:4 in. square and 4 inches thick. Vendors· have been instructed. 

to deliver their samples 1n crates on pallets, marked·as specified fn para- graph 5,4 of specification KSC-SPEC-P-0012. . . .. · 

In .addition. each vendor has been instructed to send copies of shipping notices · toM. G. Olsen, DD-m:D-1, C. L. Springfield, TG-FLD-22, and IJ, Clilutice, PRC-1217. . . to alert them of the date shipped. Freight Traffic, Sam Clymer (8G7-3NOJ 'has been requested to notify Carlos Springfield, TG-FLD-22 (067-4614) when test 

samp-les ore received so that he can direct thom to a special storage are3 main-· 

tafned by tho laboratory. . · . · . · 

. 

i 

! 

·I 

,. ..~ .. ·'· ... 

71 
NASA/TM-2010-216293R 

... '• . ..,.. ...... ·.-. 

. . ..... ' ........ "'4 .. . .. ... ' , ... . , .. . . .· .. ~.,... 

. . . \. 

Each vendor has been give the following address for shipment of samples: 

Ch1ef~ Freight Traffic. NASA 

Bldg.· 1-17-6744 

3rd St. and Avenue C 

Kennedy Space Center, FL 32899 . ATIIl: Carlos Springfield. TG-FLD-22 

I II. INSTALlATION AND R~MOVAL OF TEST Sllt-1PL(S 

· TG-SM0-1 (Wayno Parrfs/Dob Laakso) will 1n1tiate and manage the Support Request to BSI for the handling. installation. and removal of the refractory 

concrete test samples in the existing SRB flame deflector 'test fixtures. A · ... · PCN will be issued to accommodate charges. BSI 1s to provide all sarvfces. · · . ·necessary. Dolt heads and saams are to be coverad with ablative coating · · .• (D~w Corning Q3-6077) a~ specified 1n ,Owg.· 79K09S46 Sheets 1 .and 2, attached.: 

After the launch\the samples arc to be. removed and transported to an area 

designated by the laboratory. TG-SM0-1. Carlos Springfield, 867-4614. Prior 

:to removal of samples contact Carlos Springfield. Photographs are to be 

'i 

I 

' . taken of the samples prior to .their removal. 

-.IV. · SPECIM~Il- ORIEUTATION -- - --- ~ ··~ ·-- -- -· --- ·-- ••· -· ·-· - ·- -·· --· _: ~::>':~ -·· :i 

''i 

- .· ··o 

() 

' .· 

' . The test fixture on the SRD flame deflector, DWG 79K09546 will accommodate . ·six spoc1mcn samples.·· The-number of samples to bc .. submftted are,J$:1LThere- .. -.... · ·- -i 

fore. 3 launches w111 be required before the testing of all samples is .I 

accomplished. 

. ·A pattern of testing orientation in the test fixtures is as follows: 

11/1~ 

'+ 

s.s; 

* Approved for Saturn 

LAUNCH flO. 

H/W 

· ( LAUNCH tlO, 2 

for Saturn 

1.---~[ ___, 

Wahl 

N~ H/W 

Iff. .' + 

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. 0 . V. LAilORATOI~Y mASURE/IENTS OF T~ST SAMPLES 

All samples are to be tested and evaluated 1n accordance wft" KSC-SPEC-P-0012. 

Prior to installation, each test sample wi_11 be measured for an a-curate de· 

tannination of thickness using a template and a dial gage indicator mounted 

on a flat level table or surface, TG-FL0-22 to determine and perform the 

measurements, Each test sample will be photographed to record thcfr surface 

condition and texture prior to test1ng and other conditions deternlfned as 

dee.med necessary, · · ' . 

After the 1 aunch test, the specimens will bo returned to the same area where · initial measurements were made, The thickness will be remensured using the 

same techniques to determine the loss of thickness-due to general surface erosion, . ·Any_ damage or local' spa111ng or cracking w11l be noted and recorded, Each . .: . , specimen will be photographed with close-up shots of special conditions. · . :. , : . . .. ' . . 

.. ·The laboratory, TG-FLD.;22, will evaluate the rcl<~tivc performance of tha ,. ··test samples· based- upon the 'I I' f1nd1ngs · 1 n accordance with KSC-SPEC-P-0012. . . . . . 

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·. · · REFRACTORY CONCRETE, 

SPECIFICATION.FOR 

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'KSC-SPEC-P-0012. 

April 25, 1979 · 

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REFRACTORY.· CONCRETE, I' •·· • ' .... 

SPECIFICATION FOR . 

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;6 . . . . · : Approved: : . · . '· •.' 

~.<.U4K,-~ D'!rector .of Desfgn Engineering ·· 

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I JOHN F. KENNEDY, SPACE CENTER, NASA 

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' I KSC·SPEC-P-0012 

Apr1l 25. 1979 

TAill.C OF CUNTCNTS 

. SCOP[ • • • • • • Ill •• Ill ' • Ill • • ' ' Ill • • 

APPLICABLE DOCUMENTS . . .. '. .. . . . . . . . . . Governmental 

Standilrds , • 

Non-Governmental 

t I I • t t I t I 

. ............... . . . . . . . . . . . . . . . RCQU IHEMENTS . . . . . ' . . . .. . . . .. . . . . Qua11ficat ion •• .'·· ':.' -:·.-':-• _.- • .. : •••••••.••••• .. Materf als • • ·:. • • • • • . o'--:w-:_4 .• - -. • • • • • • • • • . 

Pro pert f es-·.·. • • , • • • • • • • • • _ • • •. 

..

5

F1.nenessh Modulus_-··~· • '-:.·:,;_: ~· .' • • • • •• •.:• 

trengt .,... • •:·. • •· ~ •..• ,.. • • ..• . • • • • • . . , . • • . 

Stab1l ft~.- •• ~ -_ •• -~ •••••••••••• · Rocket ··engine Exhaust Res1 stance • • • • . • • • -.. workabf11t.Y ••••••••••• _ • • • • • • •••• 

Weathering • • • • • • • • • • • • •••••• 

.QUALITY .ASSURANCE PROVISIONS . . . . . . • t • •• 

2 

'2 

2 

2 

·2 

2 

2 

2 

2 

2 

2 

Respons1bf11ty ••••••• -• • • • • • • • • • • 2 

Product Qua11ffcat ion Rcqufrerncnts • • • • • • • • • • • . 3 

Qua11flcat1onTests •••.. • • · • • • 3. 

Test Sample ••• --. ·• • , , • • • • • • • • • 3 · Re 1 nforccmcnt , • • • • • • • • • • •. • • • • • • 3 

Cover for Reinforcement. • • -, • , • • • • • • -3 

Surface F1n1sh • • • • • • • • • , • • • • • 3 

Roc kat Engf ne Exhaust Exposure • • • • • • • • • • • • • . 3 

F'i ncncss Modulus • • ·• • • • • • • • . • • • • 3 

Strength. • •• , ••••••••. , ••• , • • • 3 

Ccrtfffcate-of Conformance- • • • • • • • • • • • • • 3· 

. Test Reports • --. , • • , • • • • • • , • • • • 3 

PREPARATION FOR DEI..IVERY . . . . . . . . . . . Packag1na· ••• -· • • , •• . . . Packfng •. • • • • • • • •• 

Pa11ct1zatfon • • • •••••• 

Mark1ng • • • • • • • • • • • •••• 

Mixing and Jlpplfcat1on Instructions ••• . . . ·.,• 

.... _ 

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NO'TES ·; . . . . . . . . '.. . .. . . Int~ndcd Usc o ~ • , , 1 

Orduring Dat11 • 1 o ·, 

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'· KSC-SPEC-P-0012 

April 25- 1979 

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,!\>!IN J. 1\[NNI:llY SPACI: CCNTI:J~, MSA 

REFRACTORY CONCRETC, 

SPEC II~ I CA Tl UN FOR 

.... KSC•SP£C•P·0012 

Apr11 25, 1979 

This spec.fffcat1on has been approved by th~ OeGign Cng1nocr1ng D1r~ctorate of . the John F. Kennedy Space Center {KSC) and is mandatory for use by r.sc arid 

associated contractors. .. 

I .0 SCOPE 

Thill spccfffcatlon covers requirements for refractory concrete used for tho 

heat ilnd .blast protection of flame deflectors, and other areas of a launch 

"-----~~~-~~].:~ .. · .. ;· . . .. ~ · ... · "'•-·-~·"'~' .. :· . . . --····-··. 2.0 APPLICABLE OOCUMCNTS· ... · .. . . • .. 

0 

~· , ..... . 

The pub-.l,1c.ations ·orthe .. tsi~e~1n· ef·f~cf·-~n. the ·date ·of issuance of fnv.it~~ion for bids fonn a part of this speciffcatfon .and, where referred to thereafter 

by basic designation only, are appl icablc to the extent indicated .by the 

r·cft!rcnces thereto. In the event of difference between this spccff1catfon or 

its· accompanying drawings and the referenced· specification, thfs specification 11nd its .1ccompany1 ng drawings sha 11 oovern to the extent of such df fference.:· 

?..1. Governmental. 

?..1.1 Standards. 

Mi71tary 

Mil.·STD·129 . - Marking for Shi-pment and Storage. ·. .. . .. . . ...... - · .. ,.;..·;..-. . . . ' 

. .. · -· · (Copies of standards required by.~UJe-contr.actor in connection with specific -. · ·. procurement functions ~~~~~~IY'obtafnedffrom the procuring activity or as 

-:. ·."'Ctrrdeted b~ the. Cont~~·· f1cer.)~_:f. . _ _ . . . .· . . 

· 2.2 Non-Governmental~ · · · 

·0 

American Society for Testing and Materials (ASTM) · 

c 33 

c 39 

Concretc.Aggr$gatcs 

Compressive•Strength of Cylfndr1ca1 Concrl!te Spec1n~ens 

(1\pplication for copies should be addrcsse<l to the Amorfcdn SocfC!ty for Test· 

1ng. and Materials, 1916 Hace St., Ph1ladclphfa, Pennsylvania, 19103.) 

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·- ,• ... Apri 1 25, 1979 •· . ' . ' .,. )i.' ' •... - I 

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3.0 R[lJUIRCMEifi'S 

· 3.] (~Ill ificat1on. · The refractory concreto furnfshed under this SjJeCf fica. 

·tfon Sla e a product that has been tested and has Pdssod the qualification 

'· 

tests specified in 4.3, and has been listed or approvoo for ·listing on the ~pjJroved product~ list. · · 

3.2 _Materials. The fine aggregate shall be hard, dense, durable, clean, sharp, .Md weT! gr.lded, 

3.3 PropertfP.s. 

3.3.1 Ffnenoss Modulus. When tested 1n accordance with 4.3.2, the finenoss modulus sha It bC between 3. 75 and 2. 75. 

·• 

3.3.2 Strength. When tested in accordance with 4.3.3, refractory concrote· 

_snail develop a comprossive strength of 4500 psi (mfnfmum} at 7 days and 90. 

percent of the 7-day strength within 24 hours. If desired to develop improved properties, use of randomly dispersed steel wire fibers shall be ponnitted provided. steel fibers do not segragate and ·clog nozzles. · . 

3.4· Stabil1tl• When· maintain~ in the original unopened bag for a period of 

l year, the material sh~ll meet the r~qufrarnents of this spec:iffcatfon •. · 

3.5 ~cket Enoino Cxhclust Resistance. Test• samples ins_talloo at designated. 

areas.of the launch facility and then subject to rocket engine exhaust, shall n not c:r~ck, spall, or erode.more than l/8 inc:h~when tested fn accordanc~ with . .4.3.1.4.' Heat flux will be up to 3300 Btu/ftc;·scc; tfme of exposure wfll be ''·· . d)'proxirnately 10 seconds. . . . ' . ·;' -

3 6 Workabilftl• The refractory concrete; shall be capable of being applied 

jmewnat1c:all.Y or manually (trowel) to a uniform, smooth finish ....... ,.... · .. 

. ~ ·~-· . 

. . . 3.7 Weatherfn2• The materfa·T shaH resi~·t degradation of thermal protection characterlstJcs due to seacoast atmosphere exposure. 

4.0 QUALITY ASSURANCE PROVISIONS. 

4.1 Rospons1b1l itl• Unless othcrwf so spccf ff ed, the manufacturer· is rcspon- _sible·for tho perfOrmance of all inspection. roqufrements specified herafn, 

£xcept as otherwise specified, the manufacturer may utilize hfs own or any 

other inspection facflftfes and services acceptable to NASA. Inspection rec:- . · ords of the examinations and tests shall be kept c:ompleto and -available to. the 

Governmant. for a period of five years, unless otherwise specified in the contract or order. The Government r-eserves the rf ght to perform any of the 1 n- . spections set forth fn the specification, where such inspections arc deemed 

mants. necossary, · to en~ure supplies dnd services confonn to the prescribed rcquf re- • 

2 

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.. . r. . KSC·SPEC-P-0012 

Apr1l 25, 1979 

Ud 1 f fi Colt 1 on Rc1 ui r•ar:r~nu. To IJucorne c1 QUdl i fil!d product-, mat I!~ 

cT ::.ac:cionJ <~nd'pass the quallfic.stion tests 

4.3 Qu~lification T~sts. 

4.3.1 · ies'. Sample. 1\ test sample sha11 be required in accordance with figure 

l dnd tl1e fol1ow1ng requirernents. · 

4.3.1.1 Reinforcement. Reinforcing steel sh<lll be Bufnel Gripsteel as manuf.tctured by Keene Corpordtion. Santa Fe Springs. California or equal •. Rein·· 

rorcing steel shal·1· be free frorn rust, scale, urease, or other coating which 

rnay reduc:e the bond, 

4.3.1.2 Cover for Reinforcement. M1nirnum concreto.coverage for reinforcing 

steel from the surface exposed to the rocket engine exhaust shall not be less 

than 1-l/2 inches. · · · · 

~.3.1.3 Surface Finish, Surface finish shall be uniform and smooth. 

4.3.1 •• ·Rocket EnVfne Exhaust Exposure~· The test sample shall be installed .~t t.ht! designated aurn:h site location and exposed to a rocket engine exhaust. 

7he tes; liampl e shall be examined for confonnance to 3. 5. Install at ion and · ex ami nat 'I on of the test samples sha11 be performed by the Government. 

4.3.2 Fineness Modulus.· The fineness modulus. of the aggregate shall be : 

deterni ned 1n accordance ~ith ASTM C 33. . · · . · 

4.3.3 Stren~th~ The_compressfve s~rength shall·be detenn1ned in accordance 

with Asffi C 9. •' . . . 

4.4 Certificate of ConfonnJnce. The manufacturer shall submit a certificate 

of confonnance stabng that the rnateridl furnished is essentially identical to 

the material furnished for qu~lff1cat1on testing and complies with the 

requircntP.nts· specified herein. 

4. 5 Test Reeorts. The manufacturer shall submit a certfff ed 1 aboratory 

rupert describing the tests pcl"fonued in accordance .with 4.3.2 and 4.3.3. 

5.0 PREPARATION FOR DELIVERY 

5.1 Packaging. Unless otherwise spccff1ed, material shall be furnished in 

bags containing 100 pounds of a prernixed combination ·Of refractory aggregate fn hydraulic setting bfnder. 

5.2 Packing. Packing shil11 be in a manner which will ensure arrival at the 

designation fn satisfactory condition and be acceptable to the carrier at the 

lowest rate. · ' · 

. ,.. 

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NASA/TM-2010-216293R 

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KSC·SPEC·P·OOl Z . . Apr11 25, 1 97Y 

•I ~.:l .l• .. dll.!tizdtH,n. ~~w:1 Sl)t~Cl'fied (sue 6:2), shii>Pin~ conta1narli shd11 be 

p<llle(JZI:'d us1n~ 11ld~'dJrd wooden 1Ja11ets. · .. " I 

!:),4 Marking. In addition t.o .:~ny spuctal 111arking required by the contract, or· 

ordnr, ba~s !:hall be mJr~L'd in :.:~ccordance with MIL·STD-129. Each· bag sha11 

c.lir,play the following :nformation. 

a. Title, nwnber, and delta of this SIJOCificat1on ' . ' .. 

b. Name of the product 

c. Batch nuruber 

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d. · Manufacturer' S· name and c1ddress 

e. Weight of contents .. . . . ·. I 

f. Date of manufclcture· 

g. Toxic .. precautions 

h. :~ecessary sup!Jlernentary information to ensure safe· and proper· use ·of 

the material · · .:: 

5.5 Mixfnll and Application Instructions •. Mixing and application instructions 

shilll be 1ncluded w1th eclch sh1pmcnt. · 

6. 0 NOTCS . ..· · .,.· 

5.1 Intended use. The refractory concrete. is intended for use on .the flame .. ~eflector and other areas of a•launch complex to protect the facility from . .-· .• d~ ant heat and flame 1mpi ngcment effects of the rocket engine exhaust plume 

of d launch vehicle; 

6.2 Ordering Data. Procurement documents should specify the following: .. 

a. Title, number. and date 

. ·Of~this / spec1f1catio~ ... ~ 

b. Numbe: of 100-pound bags 

c. Cert1f1cat1on of Confonnance (see-4.4) 

d. Tc~t Reports (see 4.S) 

e. Palletfzation,·ff r~quired (see 5.~) 

NOTICE: When Government drawings. specifications. or other datcl are used for 

dny purpose other than in connection with a definitely related Government procurement operation, the United Statt~s Government. thereby incurs no responsi- bility nor any obligation whatsoever~ and the fact that the Government n~y 

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KSC-SPEC-P-0012 

April 25, 1979 

h<~'ll! fonnuliltl!d, brnfshod, or in 11ny WdY supplied the Silfd drawings, ~pi:!ciflc.H ions, or other· diltd is not to be regarded ·by imp11Cilt1on or otherwise ali in <.~ny manner 11ccns1 n!J the hul dar or any othar person or corporilt 1 on, or· convey- lny any ri(lhts or pennfssfon to manufacture, use, l)r sell any patented invan-. 

tion thdC' rnc1y In any way be related thert-to. 

CUSTODIAN: 

NI\S/1-Jolln F. J:ann~uy S)JdCl~ C~nter 

85 

PREPARING ACTIVITY: 

John F. Kennedy Space Centar 

Mechanical Design Division 

.Design Engineering Directorate 

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TYPICAL SECTION 

; . 

. 4" 

Figure 1. Test Sasplt COftffguratfon 

. .!., ____ . __ _ 

liC" 

NOTE: 

REFRACTORY 

CO~CRETE 

31"4- X 14 GAUGE 

GRID STEEL 

. REIPU'DRCISG 

ALL DIMENSIONAL TOLERA~CES 

SttALLilE: ~ 

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• II_. 

.. . /;···. .. TASK REQUEST 

• "D4T& I'-'•MITT&Q a. DC•ttt•U C01114~1r.CTtON DATK ~ I, AUTHO"I&INO OOOUMINT 

........ 4-9-76 4 .. 77 ; Nf.A 

JA~t.• DCICIIflltPTIOI!rfl . .. . : Refracto~concrete used on flame deflectors • 

....... 

I,••¥•T•M "CMO YKO P'"OMi'O" IJ•CO tNl 

N/A ( 

. 

: 

e, IIINAt.V•t• .. aou••TKCI . ·C , 

Assist fn the installation and·evaluation of refractory co~ete samples on LC-17. 

Samples to be insta11ed on carrier plates with Grfpsteel. Samples shall' be installed 

on the North side of.the umbilical tower base at a·locatfon where severe spallfng fs 

evident. 

.. Also assist in the·ident1f1cat1on of appr~ate laboratory.tests for evaluation of 

refractory concrete whfch.would be suitable for specification purposes. 

?, HICMA .. K8t 

Arrangements have been made·wfth Mr. Bob Wf 1 son, the LC-17 Fac i 11 ty Manager, for 

installation of-the test. samples. 

r 

., ... QUK."rtf.HII I, IIHON.I ID, CO~ANYI tt, MAIL. CDUICI 

~14 A~ltHGVAI.I '#,, · . Ul~ PHON&t 14, MAIL. COD"I ID,DA O.:o 

- ·M. G. Olsen~ • ~ 7-2102 DD·SE0-3 4-9-76 

FOil tAB II$E ONLif 

I .. Y"a'118AT011t TAlK NU."K"- aAM~I.· N"'fl4el" 

~oe. M<) a.(l,l ! o,.J Mm Oi~ -7 (. 

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MALFUNCTION/MATERIALS ANALYSIS SECTION. 

. ·~ATi~IALS~~NALrSJS BRA~CH 

FLUIDS·AND :ANALYSIS DIVISION ·~ . . .. 

TG-FLD-2~, ROOM 1218, O&C BUILDING 

KENNED~.SPACE CENTER, FLORIDA 32899 

JULY 29, 1981 

MMA-1918-80. 

1.0 FOREWORD 

' ... 

1.1 On 29 September, 1980, W. E. C1aut1ce, PRC-~217~ 

requested testing. of refractory concrete t.cst panels, 

the t~sting to comprise thickness measurements, color 

photographs, and visual observations. both before .and 

a.fter exposure to SRD exhaust during .STS-1 Launch.· . ' 

. 1.2 Preparations for this test had ~oen underway for some 

tfme. ·Test frames designed· by PRC had boon fnstalhd 

at LC39, · Test .panels prepared to KSC speciffcatfon had 

bee~ solicited ~rom iour vendors, Ono vendor, N~rth 

American Refractories. Company, had dcctined·t~ furnish 

samples, for reasons .. out11nod .1n their letter, :which is 

r~pr~duced· 1~· the app~nd1~. ~hree vendors had 

furnished samples as follows: 

• . 

. 1.2.1 nesianed Concretes Co. 

2 Janols Fondu Fyro WA-1 (Approv~<l for ·Saturn) 

1 panel ·Fondu ~yre WA-1 wfth wire 

1. panel F.ondu' Fyro FSC-5 

....... -::·l·.· , . 

• •• ,l 

.. ..:· 

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· ·· ... .' MMA-19-18-80 2 

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D. 

Wahl Refractory Products 

1 panel"WRPl Col or-White. 

1 panel·WRP2 .. Color-cr'ay 

· 1. panel. WRP3 Co 1 or-C.ray 

co. 

PSM Cast, HzO,,~i~er 

Cement Fondu~ ·Parry· 

Sand,·.HzO, F1bar 

. Cement Fondu· 

SOm Ba.ll Mill 

cafc1ned Fl fnt Clay 

HzO 

·.F1bor· 

1.2.3 Harbison-Walker Refractories 

· . 2 panels· Tufshot 

2 panels 1ufshot with 3.2s~·type 310 SS molt 

extracted fibers 

1.3 Shipping documents for these panels are reproduced in 

the appendix. . ' ' 

1.4 In Jhe interests of·exped1ency the support requosts_ for . 

the,hand11ng, installation. and ~omoval· of the test I 

·panels wore 1n1tiated and managed by TG-FLD-22. I 

2.0 TEST PROCEDURE 

.2.1 .. 

TEST PANEL SELECTION . .. 

The 11 test panofs supplied by the three vendors were. 

examined by. the requester. w. E. Chut1ee, PRC-1217, · 

R. J. Davis, DD·ME0-33, and C •. V. Moyers, TG-FLD-22.· 

Mr. C1aut1ce selected 6 ·panels for exposure ~urfng 

STS-1 -launch. 

89 

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NASA/TM-2010-216293R 

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' .. ' .• '·..... • J, l .t · · ,:.- MMA-1918-80 

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.2.2_:TEST PANEL IDENTIFlCATION. 

-The panels suppl1 ed· by Des1 gned Concretes bore· 1· 1 nch 

black stenc1ll~d 1dent1fy1ng m~r~s o~ the uppe~ sur-· 

face. N~ 1dent1f,ying .marks ·_were found on the: other 

panels. The sf x sel acted· panels ~-were assigned numbers 

as follows: 

. PANEL· NO· DESCRIPTION 

1.. .. Wahl. (WRP-1), Contained silvor-'colored. 

2. 

3 • .. 

4. 

5. 

· chopped _w1 re • 

. ·o~sfgncd Concretes! ."WA·1~ . 

Oes1gned Concretes,· "WA-1-W/WIRt" 

Wahl, {_WRP-3), Contained gold-colored. 

chopped wire. (WR~-3) 

Harbison-Walker, {Chopped wire}. 

6. Harbtson-Walker. · 

t.3 . PRE-EXPOSURE THICKNESS. DETERMINATIONS 

.A template to·locate 13 pri1nts on each panel (Figure 

· 1), and a .caltpcr_to determine thickness .at ~ach po~nt 

wore designed and fabricated~ The 6 panels were placed 

on .1nspect.1.on racks which afforded ~ccess to both top 

and bottom surfaces (Figure 2), and thickness at each 

of the ·predetermined points was-measured •. Photographs 

were made o.f each'panel', and they appear 1ator1nthis 

rep.ort. 

1.4 SRB-EXHAUST EXPOSURE ·· 

The panels were then installed in the two refractory 

test frames· (OWC 79K0954G) at the· bottom of the SRS · ' ' 

flame .deflector 1n the north end of the flame trench at 

LC39A:(F1gure 3). Each frame hol~~ throe panels. 'The 

holddown ~la~ps and the .areas bctwoo~ panels'were . ,., 

90 

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NASA/TM-2010-216293R 

.... •• . J_ ·;,., ·_·:· .. 

· -· MMA-1918-80· 4' ... 

' •' 

·coated wit~ Dynatherm E-320 ablative coating. The. 

panels .were- ins.talled -in numeri-ca-l sequence frorn cast 

-to west •. Figure.4 shows·a.test _f·ra~e with panels 

inst~llod~ The_pariols remained on -the test frames 

until after.STS·l launch~ 

2.5· .POST-EXPOSUR~ THICKNESS OET~RMINATIDNS 

After 1 aunch, .the panels· were photographed. They were. then 

_removed from the test frames and transported to the inspection ·racks, where the ~arne template and caliper were usod to 

make thi ck.ness determi nati,ons. The ,depths o_f ·representative 

spalled ar~~s were ·determined. using a-depth gage provt~ed 

with a df~l indicator and·a tfp with a radius of about 1/8 

f nch •. 

3.0 RESULTS 

Figures 5 through 26 show tho 1ndiv_idual panels as received, 

as installed in the_ test frames before launch,· and as thoy. 

appeared _after_ laurich. Figures 29- through 34 show indf vid~. 

ual thickness measurements before and after launch, and the 

thi.ekness. loss at each location. 

Table 1 shows tho-average thickness losses and-observations 

of surface'condftions after.~aunch. 

'· 

Based on examination .of ~he ablative coatt~g around the· 

panels after launch-. it' .appears .that tho east rack (panels 

1, _2, ~nd 3) rece1v~d more severe exposure, and that the 

.panels nearest the sides 1P~nels 1 and 6) received slightly 

less exposure tha~ the ~ther panels. on the same rack. ·Some 

variation 1n .. exp_osure fs t.o. be expected as the veh1cle· fs 

.steered. during liftoff in response to the. effects of w.f.rid at 

.the time .o·f launch·. 

',··. 

. .· ~ . 

. •. •./' . . . . . . 

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.... . '·' ·· .MMA~l918~80 · 5 

4.0 CONCLUSION.· . ' 

All of the_ panelJ lost more than·the maximum r/a~ permitted 

, by: KSC·SP.EC-P-0012 •. 

5.0. "RECOMMENDATIONS 

A .second exposure of these .panol s to tho· SRB exhaust 1 s 

recommended.·'w1th~the east to west pos1t1ons o~ the pan~ls 

changed from 1·. 2, 3- 4, 5, 6 to_4, 51 6 .. 1. 2. ·3, It is 

also ~ec~mmended'that reference ~lates w1th·a know~ 

thickness of ~blat1vc'coating bo:embcddcd flush.w1th the 

s'urface of the· ablative coating applied. to the racks. This 

_w111 permit as~cssment of the relat1ve severity of exposure 

for each panel. 

· I NVESTlGATOR :. 

C. V, MOYERS . 

. APPROVED: 

MMAS/NASA . 

92 

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MMA..;l918-80. 

Panr!l 

2. 

3. 

4. 

Average. 

·Thickness 

Loss. 

Inches 

·. 0.163 

0.256. 

0.325 

NASNTM-2010-216293R 

•.' 

6 

TABLE 1 

CONDlTION OF PANELS AFTER'·STS·l 

Average 

Thickness 

Loss. 

% of 4" 

4 .1 

6.4 

8.1-

OBSERVATIONS 

Fow _spal1 s. mostly laY"go. to. 

·0.222" deep. "Fibers~ 

exposed. 

Very few spalls.:to-0.258" 

doop. Portions of pre-exist~ 

ing crack .. covered by fused 

layor. Surface generally 

rought but uniform. Panel 

rcma1nc_d intact whon removed. 

One ipall-0.214 deep. Cracked.-

1nto six segments with some 

fused material over portions 

of cracks. Cracks measured to 

0,371" deep, with deeper 

narrow f1ssu~es •. Surface • • • I ' 

.... generally rough but uniform. 

·~:l'IPanel remained f.ntaet when 

removed • 

5 .9. 

93 

. Small, numerous spalls~ to 

0.174" deep. Gouge· at edge. 

~.37~,deop. ·orange rust 

stafn. Exposed ca~bon steel 

wlres dooply ox1d1zed. easily ,· .. broken. · 
NASA/TM-2010-216293R 

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MMA::-1918-80 

Panel 

..1!.2..:... 

5 •. ' 

6. 

Average 

Thickness 

Loss, 

Inches 

.0.238 

. 0.192. 

.·.· 

7 

'TAI3LE l 

( cont.1'nued) 

Average • 

Thickness 

Loss, 

% of 4" 

5.9 

. 4.'s· 

94 

OBSERVATIONS 

.Moderate numboi of spa11s, 

med1u~·to small, to 0.346" 

deep. "Fibers". exposed • 

Moderate. number of spalls, 

mostly large, to 0.258" deep. 

Corner inadvertently cracked· 

during removal. 

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AIEL 1, 

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100 
NASNTM-2010-216293R 

FIGttRE 7. 

PAMfl 1. IN TEST FRAME . AFT£R LAU"CH . 

101 
NASA/TM-2010-216293R 

FIGURE 8. 

'lltl 1. ON IN P!CfiO. RACK. AFTt• LAUNCH. 

102 
.------------------------------------------

NASA/TM-2010-216293R 

FIQUftE t. 

103 
NASA/TM-2010-216293R 

Fl E 10. 

PANEL 2, INSTALLED I TEST FRAME, lEFOR LAU.CM. 

104 
- -- ----------

NASA/TM -201 0-216293R 

~-1918-80 

FIGUIE 11. 

PANEL I. IN T£ T FlAM£. AF I LAUNCH. 

105 
NASNTM-2010-216293R 

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,AR£L 2 0 II 'lCTIOI lACl, AFTER LA CH. 

106 
NASA/TM -201 0-216293R 

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107 
NASA/TM-2010-216293R 

FIGURE 14. 

ALL£0 IN TEST F • I£FORE LAUNC • 

108 
NASA/TM-2010-216293R 

·1918-80 

FIGURE U. 

PAUL 3, I 

109 
NASA/TM-2010-216293R 

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AFTER LAUfiCH. 

110 
--------

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FI8UIE 17. 

111 
NASA/TM-2010-216293R 

Fl UU U. 

• 

112 
NASNTM -201 0-216293R 

tU-80 

113 
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,A L 4. GN IN 'ECTIO 

114 
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F18U E U. 

'AIEL I, AS RECIIVED. -===~----- -----~----~~~ 

115 
NASNTM-2010-216293R 

FJQURE 22. 

116 
- --- -----

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FIGUitl U. 

- ........ ......., ........... --.:.- I If TE T FltAIU • AFTER L AutiCM. _ __.. 

117 
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FIGURE 15 • 

..__.. ______ PA U I • AS RECU V£0 • __ .....;.._.....__... __ 

119 
NASA/TM-2010-216293R 

fiGURE 21. 

TALLED II TEST fRAN£. IEfotE LAUNCH. 

120 
--------------------- - ------------

NASNTM-2010-216293R 

FIGURE 27 . 

PAIIL • 

121 
NASNTM-2010-216293R 

FIGURE f1 

PANEL 6, ON INSPECTION RACK, AFTER LAUNCH. 

122 
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F!GIJR[ 29. 

THICKNESS MEASUREMENTS, PANEL I. 

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MEA~URE~ENTS,~PANEL 2. : :· '! j . ': 

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. ,., .. NORTH · ~~ERI~·~·N: .· REFRACTOR~ES CO~ · · · . 

' HANNA BUILD)NCf • EAST 14TM AND EUCLID • CLEVELAND. OHIO 441 16 • (210~ 621·6200. . 

.. ' . . ' : . ,July. 21,,' 1980 . ·. . . : " : ~ 

... -

' . . Mr·. Bill Clautico 

Planning Research Corp. · Systems Servioo Co •. 

P. o. Dox 2l2G6· 

·'· ,·. 

Kennedy S:paco Cantor, :UEl. 32815 . Ma.il Code Pn~ 1217-. _.:- · · · 

Doer Llr. Cla~~ice, ·. ·. , .. -. 

Co~1rming Our :phone conversation tlufl. dc.te, thio lottor .will . confirm our company docioion·not· to !urniOh toot aam:pleo of re!-

rnotory. ponoreto i'or·:poeeible uee·at laWlCh ai'tes • . ·. 

,. , I •. 

As we discussed, ~ of· the• 'spe'cifioo.tiono outlined in KSC-SPEC-P- .. ·. 0012 wero toote not normally :performed on rogulo.r rofrnotoey llrO~ . .· uots, but ra.ther were ·.teoto U3Sd on· rogulo.r portland oemont •. Some.,·::.::; .. · of tho finenoBe.recwoments. would.{la.ve. limited us in tho typo of, ... :··. 

product to bo rooommondod and lilomo of·tho othor roquiremento raised · I qu~_otiono not pr.ev1oucly onoountero(l.. . · . · .. I 

We regret ~1ut we will be unable to furnish tho toot camploa you roquire and hope · you will bo o.blo to o btllin. onough. onmplee to ru.n, 

tho toot you ha.ve. 1ndioo.ted •. : · : ·. . . . · .. . . 

. .. . Porhn:po in tho fUture it:'you·.:have·::o.ther ·:neodo::·we··wiii ~be eble. to . . work with you o.t tho.t time •. If wo oc.n be of oorvico in any ·way,: .. · ··:· .. · · bo curo to oall on uo. . .. · · · · 

S1nceroly, · ·. · · 

f;.r; ~ L I .·. 

G. Tro.ot~""'· . 84 Hatch.inoha Road Irt~1nos 01 ty, Flo.. 33644 

813.:.439-4519 . . 

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CO: R. A. LUnd · J. ::.I. so·o.nilon 

E. s. Chrzan 

114 HATCHIN(HA RD. ' 

HAIN($ ClTY, FLORIDA 331U ' . TILI!PHONI .. --,; 

O"'C! I 113.~39·~519 . : HOMEr 813·439-3~68 

. nooearoh. Dept • . File · 

' · . 

G. TRU!Tr LANFORD • , Sol .. R.,......,,cothoe . 

gtl/o 

. ··t) NORTH AMERICAN REPRACTORiaa ·~c, · . . .. [Elba)_, . · · . ; 

130 
. -r-:··~."·,·-- ···--'7"~. ' . ' . ' '' ' • .. •• •.. ,., ,. ,·~ ' : ,., •'" • • •, •' •••: I, • 'l , . .:.. ' '\'1'1 j o' I I I' 

"'IG,HT SILL OF LADING-SHORT FORM..:.. Original-Not Nogotlable •. I, . ' •,' ' ' ',' 

......... 

J. I .·. ,f ·'. :: 

:;; 

CHIEf ~RT~ .TRAFFIC,;~ASA~· eLDG r~7-6744 . . . .•. 

3RO STREET'ANO'.AV~·C 

KENNEDY SPC .. CNT FL 32899 

Attn: ·ca~los Springfield 

131 

NASA/TM-2010-216293R 

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NASA/TM-2010-216293R 

', ·.&.o .. Ozolo•N--·, . . •.· 

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DESCRIPTION 

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Phone (419) 334-2658 · 

46 

REFRACTORY PRODUCTS COMPANY 

P. 0. Box 430; Green Spring• Road, Fremont. Ohio 43420 

January 7. 1981 

Mr. Mel Olsen 

J •. F. 

Kennedy Space Center. Fla. 32899 

00-MED-1 

Dear Mr. Olsen: 

The follaw1ng·1nformat1on concerns our telephone conversation of January 7. 1981. . -~~ 

Panel' II . Constituent· 

WRP 1 Color-White PSM· Cast 

H2o 

Fiber . ' 

------~------------·M·----------------------------------------------

WRP 2 co.l or-Gray .. 

WR~ 3 Color-Gr~y · 

.. 

Cement Fondu 

Parry Sand 

H20 

Fiber 

Cement Fondu 

50m Ball M111 

Calcined Flfnt Clay 

3/F Flint Clay .. H20 

Fi6er 

·If yo~ have any questions concerning .this 1nformat1on. please' don't hesitate to 

let me know. 

'l 

:I 

.I 

1 

:I j 

,A Sincerely, . 

\.WAHL.~ ~F~ F~. PRODUCTS COMPANY 

. ~J\ \S.~~ . ':J . \ 

Danfel H. 'Lease /•· ,Uf'. i'/ 

I O: ·, ' 

I 

1 

I I ,o ,' 

CASTABLES PLASTICS MORTARS REPRACTORV ENGINEERING 

133 
NASNTM -201 0-216293R 

--· .. ~-... -............... -.. -..---~·----···-·••'"-·••·•--·-·•·&-- VIIIIVII ...... II11WV..._~--·_.,..,~ .............. ........ : ..... ·' - •· ...... ·' • •.~ ...... . 

r-''r·; .. ::: .. the w~oo~ ~~~~ry prodv~;s··~o~;:;~· R. p 47 1 ., • • HIGH ·TEMPERA 'tURf: CEMENTS ·.":"•· • • '• Pl.ASTlC•.I'IRE'6RlCK · 

I '• 

'I 

ROUTI 

767 S. R. 19 South, 

NASA 

Chief, Freight Traffic 

Bldg. M7-G744 

3rd. St. &·Ave. C. 

', 

' 

P. 0. Box 430 · 

I ''• 

·Kennedy Space·Center, Florida 3Z899 

Premont, Ohio 43420 

Same 

Bl.dg. M7-G744 · , Att: · Carlos· Springfield 3Rd.· St. & Ave. C. 

Kennedy Space Center, F1or1d~ 

~~~o~.~~~~ .. ~a--------~~------------~~~~~~--~~~~..----

de 'TG-FLD-22 2·12·80 

,. 

' 

.... 

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3 

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TERMS• 

Test Panels (approx. 190# ea.) 570# ·,•' 

l Pallet 

· . 

·TEST : NO CHARGE 

( 

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~ r.: ·: .. ri': 

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NASA/TM-2010-216293R 

APPENDIX B. MTS-505-81, STS-2 

135 
NASA/TM-2010-216293R 

'· NASA 

MATERIALS TESTING SECTION 

MATERIALS ANALYSIS BRANCH 

FLUIDS AND ANALYSIS DlVISlON 

TG-FLD-22, ROOM 1218, O&C BUILDING 

KENNEDY SPACE CENTER, FLORIDA 32899 

MARCH 1, 1982 

MTS 505-81 

SUBJECT: Exposure Test of Refractory Concrete Test Panels to 

Solid Rocket Booster (SRB) Exhaust During STS-2 Launch 

RELATED DOCUMENTATION: MMA-1918-80, July 29, 1981 

1. 0 FOREWORD 

1.1 On 3 September, 1981, Gary Kurtz, PRC-1217, requested 

testing of refractory concrete test panels, the testing 

to comprise thickness measurements, color photographs, 

and visual observations both before and after exposure 

to SRB exhaust during the launch ~f STS-2. 

1.2 This test is a continuation of the testing reported in 

MMA 1918-80, dated July 29, 1981. The test panels 

provided by vendors and the test fixtures are described 

in that report, which covers the exposure of 6 panels 

to SRB exhaust during STS-1. 

·----··-·-·--·-··--. ..-...... ·.-. ··-~--~·--·----------·· -· 

136 
NASNTM-2010-216293R 

MTS 505-81 2 

.~ 2. 0 TEST PROCEDURES 

~·. 

2.1 TEST PANEL SELECTION AND IDENTIFICATION 

In Test Report MMA 1918-80 a second exposure of the 

initial panels to the SRB exhaust was recommended, with 

the east-to-west positions of the panels changed from 

1, 2, 3 - 4, 5, 6 to 4, 5, 6 - 1, 2, 3. It was also 

recommended that reference plates with a known thickness of ablative coating be embedded flush with the 

\ surface of the ablative coating applied to the racks. 

This was intended to permit assessment of the relative 

severity of exposure for each panel. At the request of 

the vendor, Wahl Refractory Products Company, test 

panel 1 was removed from the test. It contained 

chopped stainless steel wire whi~h increased its cost 

with no apparent improvement in performance. This 

panel was replaced by a previous unexposed panel. The 

panels selected with their east-to-west exposure 

positions are as follows: 

Exposure 

Position 

East-to-West 

1 

2 

3 

4 

5 

6 

Panel 

NO. 

4 

5 

6 

7 

2 

3 

Wahl (WRP-3), Contained carbon 

steel wire 

Harbison-Walker Tufshot, 

Contained "fibers" (310 ssl 

Harbison-Walker Tufshot. 

Wahl (WRP-2) (Previously 

unexposed) 

Designed Concretes, Fondu Fyre 

WA-1 

Designed Concretes, Fondu Fyre 

WA-1 w/W ire 

137 
NASNTM-2010-216293R 

MTS 505-81 3 

2.2 PRE-EXPOSURE THICKNESS DETERMINATIONS 

The 6 panels were placed on inspection racks which 

afforded access to both top and bottom surfaces, and 

thickness at each of the predetermined points was 

measured, using the template and caliper employed in 

previous measurements. Photographs· of each previously 

~xposed panel, made at that time, appear in 

MMA-1918-80. Figure 10 shows panel 7 before exposure. 

2.3 SRB EXHAUST EXPOSURE 

The panels were then installed in the two refractory 

test frames (DWG 79K09546) at the bottom of the SRB 

flame deflector in the north end of the flame trench at 

LC39A. Each frame holds three panels. The holddown 

clamps and the areas between panels were coated with 

Dynatherm E-320 ablative coating. The panels were 

installed in the positions indicated in paragraph 2.1. 

Reference panels were embedded flush with the surface 

of the ablative coating in areas between test panels. 

The panels were then exposed to SRB exhaust during 

STS-2 launch. 

2.4 POST-EXPOSURE THICKNESS DETERMINATIONS 

After launch, the panels were examine~ and photo~ 

graphed. It was observ~d that the steel reinforcing 

grid of t~st panel 6 in position 3 was completely 

exposed. Later, when photographs were being made, it 

was found that position 3 was empty (see photograph, 

Figure 9), It is conjectured that the remaining 

portions of panel 6 may have been considered loose . 

debris by the cleanup crew and removed inadvertently 

during cleanup of the flame trench. It was also 

observed that none of the reference plates remained in 

place after the launch. 

·-~ --·----.~ .......... - --~,---·--·· 

138 
\ 

NASA/TM-2010-216293R 

MTS 505-81 4 

~. The panels were then removed from the test frames and 

transported to the inspection racks, where the same 

template and caliper were used to make thickness determinations. The depths of representative spalled areas 

were determined, using a depth gage provided with a 

dial indicator and a tip with a radius of about 1/8 

inch. 

3.0 RESULTS 

Figures 1 through 12 show the individual panels in the test 

frames after launch, and on inspection racks after launch, 

as well as panel 7 as received. Figures 13 through 17 show 

individual thickness measurements before and after STS-1 and 

STS-2 launches, and the thickness losses at each location. 

Table 1 shows the average thickness losses and observations 

of surface conditions after launch. 

3.1 All surviving previously-exposed panels showed surprisingly lower losses during their seco~d exposure 

than during their first. Indeed, in a number of locations, there was an actual increase in thickness during 

STS-2 launch. This is attributed to exhaust residue 

buildup. Tests performed after STS-1 launch indicated 

that material on the surfaces of panels after exposure 

was probably SRB e~haust product. This work was not 

completed before the publication of MMA 1918-80. A 

report of the analysis is therefore included in this 

report as Appendix 1. 

3.2 Panel 7, which was exposed for the first time during 

STS-2 launch, was in the worst condition of the 

panels surviving this launch. One corner was badly 

damaged, and large, deep.pits covered a large part of 

its surface. Nevertheless, thickness measurements in 

some predetermined locations showed losses comparable 

to those of the other panels during their first 

exposure. 

139 
NASA/TM-2010-216293R 

MTS 505-81 5 

4.0 CONCLUSIONS 

4.1 The effect of previous launch exposure and differences 

i~ lauhch conditions could be factors .in causing the 

difference in material lo~ses du~ing the first and 

second launch. For example, about four times as much 

water was consumed during the second lau~ch as during 

the first launch because of changes in the sound 

suppression' system. 

4.2 Reference panels for exposure severity assessment must 

be securely fastened. Panels embedded in ablative 

coating as the only securing means will not survive a 

launch. 

4.3 Pan~l 7 fs badly deteriorated, and further exposure is· .. not warranted. 

4.4 The following panels are now on hand: 

Number of 

Panel Launch 

No. 

1 

2 

3 

4 

5 

Exposures Description 

1 (STS-1) Wahl WRP-.1. Contained "fibers" 

(stiinle~s steel wire) 

2 Designed Concretes, Fondu Fyre WA-1 

2 Designed Concretes, Fondu Fyre WA-1 

w/W:i re 

2 Wahl WRP-3. Coritained carbon steel 

wire 

2 Harbison-Walker Tufshot. Contained 

"fibers" (310 ss) 

7 1 (STS-2) Wahl WRP-2 

Undesig- 0 Designed Con~retes, Fondu Fyre WA-1 

nated (same as Panel 2) II 

II 

0 

o· 

0 

Designed Concrete~, Fondu Fyre FSC-5 

H~rbison-Walker Tufshot (same as 

Panel 6) 

·Harbison-Walker Tufshot with 

"fibers" (310 ss) (same as Panel 5) 

140 
NASA/TM-2010-216293R 

MTS 505-81 

~- 5. 0 RECOMMENDATIONS 

5.1 Six panels from the above list should be selected for 

exposure to STS-3 launch exhaust. Panel 7 should not 

be considered. 

6 

5.2 If time permits, a method of securing reference panels 

for assessing severity of exposure should be devised. 

INVESTIGATOR: C.l/· (YII")I?'1 

C. V. MOYERS 

C. L. SPRINGF 

141 
NASA/TM-2010-216293R 

,-..__ 

MTS 505-81 7 

Panel 

No. 

2 

3 

4 

5 

6 

7 

TABLE 1 

CONDITION OF PANELS AFTER STS-2 LAUNCH 

Average 

Thickness 

Loss, 

Inches 

0.087 

0.073 

0.003 

0.063 

0.259 

Observations 

Coated with residue, cracks healed, pits 

to 0.311" deep, generally uniform 

appearance. 

Coated with residue, cracks healed, pits 

to 0.201" deep, generally uniform 

appearance. 

Coated with residue, few pits to 0.278" 

deep, generally uniform appearance. 

Heavy coating of residue, but many 

uncoated'irregular pits to 0.656" deep. 

Cracks healed in places, mostly open. 

Steel grid completely exposed after 

launch. Remains of panel not recovered. 

Severely eroded, pits to > 1.00" deep. 

------·--· ... ---~--- -- -----· ~--~---··-- ---- . 

142 
------------------------

NASA/TM-2010-216293R 

MTS-505-81 8 

FIGURE 1 

PANEL 2 IN TEST FRAME, POSITION 5, AFTER STS-2 LAUNCH. DESIGNED 

CONCRETES FONOU FYRE WA-1. 

143 
NASA/TM-2010-216293R 

FIGURE 2 

PANEL 2 IN INSPECTION RACK AFTER STS-2 LAUNCH 

144 
MTS-505-81 

NASA/TM-2010-216293R 

FIGURE 3 

PANEL 3 IN TEST FRAME, POSITION 6, AFTER STS-2 LAUNCH 

DESIGNED CONCRETES FONDU FYRE WA-1, W/WIRE 

145 

10 
NASA/TM-2010-216293R 

FIGURE 4 

PANEL 3 IN INSPECTION RACK AFTER STS-2 LAUNCH 

146 
- --------

NASA/TM-2010-216293R 

FIGURE 5 

PANEL 4 IN TEST FRAME, POSITION 1, AFTER STS-2 LAUNCH 

WAHL WRP-3 WITH CARBON STEEL WIRE. 

147 
NASNTM-2010-216293R 

MTS-505-81 

FIGURE 6 

PANEL 4 IN INSPECTION RACK AFTER STS-2 LAUNCH. 

148 
NASA/TM-2010-216293R 

MTS-505-81 

FIGURE 7 

PANEL 5 IN TEST FRAME, POSITION 2, AFTER STS-2 LAUNCH. 

HARBISON-WALKER TUFSHOT WITH "FIBERS" (310SS) 

149 
NASA/TM-2010-216293R 

MTS-505-81 15 

. "-.. 

• . . 

.-

• .. 

• 

FIGURE 8 

PANEL 5 IN INSPECTION RACK AFTER STS-2 LAUNCH 

150 
NASA/TM-2010-216293R 

MTS-505-81 16 

FIGURE 9 

TEST FRAME, POSITION 3, AFTER STS-2 LAUNCH. THE REMAINS OF PANEL 

6, HARBISON-WALKER TUFSHOT, HAD BEEN REMOVED BEFORE PANEL 

RETRIEVAL. 

151 
NASNTM-2010-216293R 

MTS-505-81 l7 

FIGURE 10 

PANEL 7, WAHL WRP-2, BEFORE EXPOSURE TO STS-2 LAUNCH 

152 
NASA/TM-2010-216293R 

MTS-505-81 

FIGURE 11 

PANEL 7 IN TEST FRAME, POSITION 4, AFTER STS-2 LAUNCH 

153 
NASA/TM-2010-216293R 

MTS-505-81 

FIGURE 12 

PANEL 7 IN INSPECTION RACK AFTER STS-2 LAUNCH 

154 
NASA/TM -201 0-216293R 

MTS-505-81 

20 

STS-1 STS-2 STS-1 STS-2 STS-1 STS-2 

4.105 3.920 4.120 3.910 4.135 3.920 

0 3.920 3 828 0 3.910 3_,808 3.920 0 3.833 

11 12 o:·102 13 0.185 0.092 0.210 0.215 0.087 

- STS-1 STS-2 STS-1 srs .. 2 

4.090 3.850 4.100 3.780 

·o 3.850 3.748 0 3.780 3.728 

9 10 0.240 0.102 0.320 0.052 

STS-1 STS-2 STS-1 STS-2 STS-1 ' STS-2 

4.145 3.850 4.090 3.820 4.120 3.800 

0 3.850 3.728 0 3.820 3.693 3.800 0 3.788 

6 7 8 0.295 0.122 0.270 0.127 0.320 0.012 

. 

STS-1 STS-2 STS-1 STS-2 

4.100 3.840 4.106 . 3.860 

: 0 3.840 3. 778 0 3.860 3.74~ 

4 0.260 0.062 5 0.246 0.112 

' 

"", STS,..1 STS-2 STS,..1 . STS-2 .STS-1 STS-2 

. 4.166 3.940 4.120 3.860 4.175 .. 

3.900 

0 3.940 3.868 0 3.860 3.798 .·, 3.900 0 3.768 

1 0.226 0.072 2 0.260 0.062 3 0.275 0.132 

'· 

. 

AVERAGE THICKNESS LOSS 

STS-1 0.256 

~-"i-2 0.087 

,OTAL 0.343 FIGURE 13 

THICKNESS MEASUREMENTS, PANEL 2-, DESIGN CONCRETES FONDU FYRE WA-1 

155 
NASNTM-2010-216293R 

MTS- 5.05-81 

STS-1 

4.045 

0 3.790 

11 0.255 

STS-1 

4.040 

0 3. 710 

6 0.330 

STS-1 

4.080 

0 3.780 1 0.300 

AVERAGE THICKNESS LOSS 

STS-1 0.325 

,r-·s-2 o.073 

oOTAL 0.398 

STS"-2 

3.790 

3.748 

p.042 

·o 

9 

STS-2 

3.710 

3.668 

0.042 

: 0 

4 

STS-2 

3.780 

3.728 

0.052 

STS-1 

14.125 

0 3.850 

12 0.275 

. STS-1 STS-2 

4.065 3.760. 

3.760 3. 728 

0.305 0.032 

STS-1 

4.055 

0 3.700 

7 0.355 

. STS-1 STS-2 

4.030 3.680 

3.680 3.458 

0.350 0.222 

STS.-1 

4.045 

0 3.700 2 0.345 

FIGURE 14 

21 

STS-2 STS-1 STS-2 

3.850 4.090 3.860 

3.848 3.860 0 3.778 . 

o:~·oo2 0.230 13 0.082 

STS-1 STS.-2 

4.100 3.770 

0 3. 770 3:688 

10 0.330 0.082 

STS-2 STS-1 ' STS-2. 

3;700 4.115 3.650 

. 3.598 3.650 0 3.558 

0.102 0.465 8 0.092 

STS-1 STS-2 

4.030 3.690 

0 3.690 3.578 

5 0.340 0.112 .• 

·\-. 

STS-2 STS-1 STS-2 

~ 3.700 3 .. 720. 

3.708 

.·• ··. ~ 3.628 

-0.002 . 0.092 

'· 

. 

THICKNESS MEASUREMENTS, PANE,L 3, DESIGN CONCRETES FONDU FYRE WA-1 W/WIRE 

156 
MTS-505-81 

STS-1. STS-2 

4.502 3.970 

0 3.970 3.948 

11 0.235 0.022 

·o 

STS-1 STS-2 

4.140 3.900 

0 3.900 3.798 

6 0.240 0.102 

: 

STS-1 STS-2 

4.190 .3. 980 

0 3.980 3.978 

1 0.210 0.002 

AVERAGE THICKNESS LOSS 

STS-1 0. 235 

(·'s-2 o.oo3 

.• OTAL ·0.238 

9 

0 

4 

STS-1 

4. 135 

0 3.960 

12 0.175 

- STS-1 STS-2 

4.110 3.840 

3.840 3.798 

0.270 0.042 

STS-1 

4.055 

0 3.790 

7 0.265 

STS-1 STS-2 

4.145 3.890 

3.890 3.898 

0.255 0.008 

STS~1 

4.205 

0 3.990 

2 0.215 

FIGURE 15 

NASA/TM-2010-216293R 

22 

STS-2 STS-1 SiS-2 

3.960 4.120 3,870 

3.938 3.870 0 3.908 

0': 022 0.250 13 0.038 

STS~1 STS .. 2 

4.155 3.840 

0 3.840 3.858 10 

0.315 -0.018 

STS-2 STS-1 : STS-2 

3.790 4.175 3.970 

3. 778 3.970 0 4.010 

0.012 0.205 8 ~. 040 

STS~l STS~2 

4.205 3.990 

0 3.990 4.018 

5 0.215 -0.028 .•. 

~ 

STS-2 .STS-1 STS-2 

3.990 4.205 4.000 

4.000 ··- 4.000 0 4. 020 . 3 

0.010 0.205 .{).020 ' 

THICKNESS MEA,SUREMENTS, PANEL 4, WAHL WRP-3 

157 
NASA/TM-2010-216293R 

. MTS-505-81 

STS-1 STS-2 

4.190 3.980 

0 3.980 3.938 

11 0.210 .0.042 

·o 

STS-1 STS-2 

4.250 4.010 

0 4.010 3.918 

6 0.240 0.092 

: 

STS-1 STS-2 

4.175 3.870 

0 3.870 3.968 

1 0.305 -0.098 

AVERAGE THICKNESS LOSS 

STS-1 0.238 

.;-'S-2 0.063 

a~JTAL 0.301 

9 

0 

4 

STS-1 

4.255 

0 4,050 l2 0,205 

·- STS-1 STS-2 

4.180 3.950 

3.950 3.898 

0.230 0.050 

STS-1 

4.330 

0 4.030 

7 0.300 

·STS-1 STS-2. 

4.155 3.910 

3.910 3.868 . 

0.245 0042 

STS-1 

4,040 

0 3,850 

2 0.190 

FIGURE 16 

23 

STS-2. STS-1 STS-2 

4,050 4.210 4.040 

3,990 4.040 0 3.938 

rf:o6o 0.170 13 0~ 102: 

STS-1 srs .. 2 · 

4.480 4 280 

0 4.280 0.200 10 

0.200 0.080 

STS-2 STS-1 ' STS-2 

4.030 4.270 4.000 

4.040 4.000 0 3.988 

8 -0.010 0.270 0.012 

STS-1 STS-2 . 

4.330 4.080. 

0 4.080 3.928 

5 0.250 0.152 -

\,. 

STS-2 STS-1 : STS-2 

3.850 4.225 .. 3;950 

3.598 - 3.950 0 3.908 . 3 0.252 0.275 0.042 : 

. 

THICKNESS ~1EASUREMENTS, PANEL 5, HARBISON-WALKER TUFSHOT \41TH "FIBERS1' · 

158 
NASA/TM-2010-216293R 

MTS-505-81 

24 

STS-1 STS-2 STS-1 STS-2 STS-1 

11 0~ 12 0~ _,:.._ ~~ 

0 

6 

0 

1 

AVERAGE THICKNESS LOSS 

STS-1 

~S-2 0.259 

.uTAL 0.259 

- STS-1 STS.-2 

m am ·o 

9 . 

STS-2 STS-1 STS-2 STS•l ~ STS-2 

4.050 

~·~ 3.960 

3.678 0 3 898 7 

0.372 0.062 

STS-2 STS-2 

4,010 3.990 

0 3,498 3.733 

4 0.512 0.257 

.... STS-2 STS;..l STS-2 

3.990 4,010 

··~·~ 3.91Q 

3.618 0 3.748 3~838 . 2 0.372 0.262 0.072 . . 

FIGURE 17 · 

THICKNESS ;MEASUREMENTS, PANEL 7, WAHL WRP-2 

159 
. NASA/TM-2010-216293R 

. rMICROCHEMICAL ANALYSIS SECTION TG~FLD-21, Room 1274, O&C Building NASA/KSC 

Sept. 16, 1981 

SUBJECT: Analysis·of.Refractory Concrete Test Panels 

LABORATORY REQUEST NO: MAS-310-81 

1 . 0 Foreword 

1.1 Requester: C .. V. Moyers/NASA/TG-FLD-22 (867-4614) 

APPENDIX I 

1.2 Requester's Sample Description: Material removed from refractory 

concrete test panels after exposure to SRB Exhaust During STS-1 

Launch. Panel 1 -Surface layer and base material. Panel 2-

Fused material and base material. 

1.3 Requested: Are blackened surface· and fused material refractory 

concrete or SRB exhaust products? 

2.0 Chemical Analysis and Results 

2.1 The samples were analyzed by X-ray fluorescence spectrometry for 

qualitative elemental composition and by X-ray powder diffractometry for phase identification . 

2.2 The results of both types of analysis are presented in Table I. 

3.0 Conclusions · 

.Alpha aluminum oxide (corundum), the phase known to be formed in solid 

rocket booster exhaust product, was found.in both the blackened surface 

and the fused material, but not in either Panel 1 or 2 base material. · It therefore appears likely that the b 1 ackened surface and fused materia 1 

are SRB exhaust products and not refractory concrete. 

160 

Chemist: JJ IJ . L._...zr- H. D. Bennett . 

.. ,,.,..,,~ nes 
-

0'1 -

Sample 

Panel 1 - Blackened 

Surface 

Panel 1 - Base 

Material 

Panel 2 - Fused 

Material, Downstream 

Edge 

Panel 2 - Base 

Material 

MAS-' ')81 

TABLJ;. i 

X-Ray Fluorescence 

MajorAl, Ca 

Minor Fe, Ti 

Trace Zr, Zn, K, Cl, S, Si, Ni 

Major Al, Ca, Ti 

Minor Fe, K, Si, Zn 

Trace Ni, S, Zr 

Major Al, Ca, Ti, Fe 

Minor Zn, K, Si 

·Trace Ni, Cl, S 

Major Ca, Ti, Fe 

Minor K, Si 

Trace Zn, S, Cl, Ni, Zr 

X-Ray Diffraction 

alpha Aluminum Oxide, «- Al203 (Corundum) 

Silicon Oxide, Si02 (Cristobalite) 

Aluminum Silicate, 3Al203 • 2Si02 

(Mullite) · Silicon Oxide, Si~ (Quartz) Calcium Carbonate, CaCOJ (Aragonite) Aluminum Oxide Hydrate, Al203 • 3H20 

(Gibbsite) 

alpha Aluminum Oxide,«- Al203 

(Corundum) 

Titanium oxide, Ti02 (Rutile) 

Iron (II, III) Oxide, Fe304 (Magnetite) Calcium Magnesium meta Silicate, 

CaMg(Si03)2, (Diopside) Additional unidentified phases, pro- bably silicates 

... 

)> 

""0 

"'0 

I"T'' 

z 

0 

...... 

X 

...... 

z 

> r/1 

~ ~ I 

N 

0 -

0 

I 

N -

0'1 

N 

\0 

w 

:;:tJ 
NASA/TM-2010-216293R 

.---· 

DISTRIBUTION LIST 

SYMBOL/NAME NUMBER OF .COPIEs· 

PRC-1217/KURTZ 5 

ZK86/VAN DUSEN 1 

PR0-1214/WALKER 1 

DD-MED-1/0LSEN 1 

SI~PR0-14/N. PERRY 1 .... -.. .. 

SF~ENG/ROBERT- MILLER 1 

SF~ENG-3/R. GILLETT 1 

TG~FLD~2 1 

TG~FLD-22 ALL THAT ARE LEFT 

162 
NASA/TM-2010-216293R 

.. TASK REQUEST 

1. DATE SUBMITTED. 2, DESIRED COMPLETION DATE 8, AUTHORIZING DOCUMENT 

3 Sept 81 One Month after Launch of 

STS-2 

SAMPLE DESCRIPTION: 

Refractory Concrete Test Samples:· 

1. Design Concrete Company '(a) Fondu Fyre WA-1 (Approved for Saturn) '"' 2 Test Samples (b) Fondu· Fyre WA-1 (Experimental) - 2 Test Samples 

2. Wahl Refractory Products Company (a) H. T. Bond Mortar (Experimental) - 3 Test Samples 

3. Harbison-Walker Refractories 

(a) Refractory Concrete -------------- 2 Test Samples (b) Refractory Concrete (Stainless Steel Fibers) - 2 Test Samples 

8. SYSTEM REMO VEO FROM/OR USED IN: 

Samples for test. - Refractory Concrete 

Location to be SRB Flame Deflector, 79K09546. 

e, ANALYSIS REQUESTED: 

SEE "REFRACTORY MATERIALS TEST PLAN" - attached 

The materials are on hand. Six panels are to be exposed to SRB Exhaust during STS-2. 

Testing is to be initiated by taking thickness measurements and color photographs 

of the initial condition of the surfaces . ... 

After exposure to SRB exhaust during STS-2, the samples are to be removed and 

returned to the laboratory for further thickness measurements and a determination 

of surface erosion. The final s·urface condition such as spalling will be recorded 

and color photographs taken for comparison with initial surface conditions. 

An evaluation of performance and acceptance or rejection recommendations will be 

made by formal report. 

-

7. REMARKS: 

8. REQUESTER; 

PRC 1217 ~"""i ~ g, PHONE: 10, COMPANY: 11. MAIL CODE: 

G. KURTZ, 867-3407 PRC PRC-1217 

~.NASA AP~VAJ ~~ ' ·:;;:;:z 14. MAIL CODE: 18, DATE: 

. , -~ 

R. . DAVIS -"'::'""~ OLSetl DD-MED-33 ~-'-t' • FOR LAB USE ONLY 

INVESTIGATOR TASK NUMBER SAMPLE NUMBER 

~ _p--rs 5:" 5- 1?' I KSC FORM 22 81 {REV, 1/701 

163 
NASA/TM-2010-216293R 

l:lFRACTORY COI~CRETE 1·!A-:-E;.;IALS TEST PLAH 

1. fJif\ TERIALS FCt\ TESTIIlG ( KSC-SPEC-P-0012) 

~ ' 

1. Design Concretes Company 

a. Fondu Fyre WA-1*- 2 Test Samples 

b. Fondu Fyre I·IA-1 (Experimental) - 2 Test Samples Delivery date July 21, 1980 

* Approved for Saturn 

2. Wahl Refractory Prod~cts Company 

a. H. T. Bond l~ortar (Exp!)rimental) - 3 Test Samples 

The 3 test samples are at KSC. 

The vendor delivered them on Feb. 12, 1980 

They are stored by TG-FLD-22. 

3. Harbison - Walker Refractories 

a. Refractory Contrete - 2 Test Samples 

b. Refractory Concrete - 2 Test Samples (Stainless Steel FifJers) Delivery Date June 20, 1980 

' L.ETE]) @ 1lorth American Refractories Company 

a. Narco Cast 60 - 2 Test Samples . b. Narco Tab- 2.Test Samples Delivery date Aug. l, 1980 

I I. TEST SAMPLE DELIVERY, HAilDLillG AND STORAGE 

The refractory concrete test samples as specified in KSC-SPEC-P-0012, Figure l. · are massive and require special handling. They weigh approximately 200 lbs. 

each and are 2 Ft-4 in. square and 4 inches thick. Vendors have been instructed 

to deliver their samples in crates on pallets, marked as specified in~=~=­ graph 5.4 of specification KSC-SPEC-P~0012. 

In addition, each vendor has been instructed to send copies of shipping notices 

to H. G. Olsen, DD-MED-1, C. L. Springfield, TG-FLD-22, and U. Clautice, PRC-1217 

to alert them of the date shipped. Freight Traffic, Sam Clymer (867-32.:2~ has been requested to notify Carlos Springfi1~ld, TG-FLD-22 (867-4614) h~e~ test 

san;ples are received so that he can direct them to a special storage are3 ~·ain­ ta ined by the laboratory. 

164 
NASA/TM-2010-216293R 

,~ Each vendor has been give the following address for shipl"ent of samples: 

r'·· 

Chief, Freight Traffic, NASA 

Bldg. 1·17-6744 

3rd St. and Avenue C 

Kennedy Space Center, FL 32899 

ATTN: Carlos Springfield, TG-FLD-22 

I I I. IrJSTALLATION AND REMOVAL OF TEST SAI-1PLES 

TG-SMD-1 (WaYne Parris/Bob Laakso) will initiate and n:anage the Support Request to BSI for the handling, installation, and removal of the refractory 

concrete test samples in the existing SRB flame deflector test fixtures. A 

PCN will be issued to accommodate charges. BSI is to provide all services 

necessary. Bolt heads and seams are to be covered with ablative coating 

(Dow Corning Q3-6077) as specified in Dwg. 79K09546 Sheets 1 and 2, attached. 

After the launch the samples are to be removed and transported to an area 

designated by the laboratory, TG-SMD-1, Carlos Springfield, 867-4614. Prior· 

to rer.1oval of samples contact Carlos Springfield. Photogrqphs are to be 

·taken of the samples prior to their removal. 

IV. SPECI14W ORIEIHATION 

The test fixture on the SRB flame deflector, DWG 79K09546 will· accommodate 

six specimen samples. The number of samples to be submitted are .]..8"/1 Therefore, 3 launches will be required before the testing of all sampies is 

accomplished. · 

A pattern of testing orientation in the test fixtures is ~s follows: 

Wa11T Fondu 

HT Fire ~\Ala.~\ 

Bond WA-1 ff*P7 ~1ortar * 

F'<nl»•' r-Ng:. H/W 11/H 

+ wA~""' (E'1i' s.s. 

I * Approved for Saturn 

LAUNCH llO. 

Fan u .-Wahl 

H/W 

~ rw Fyre . m--

WA-1 Bem:t-

* llal'ta!' 

l' *Approved for Saturn I LAUNCH NO. 2 

H/vl I I 1 1 iiY Narco + 

' Eond TAB s.s. 

: ::ortar! l I l I 

.L--- . L._ __ _ _ __j 

LAUNCH 116. 3 

165 
NASA/TM-2010-216293R 

I . 

. ~· V. LABORATOI~Y ~lli\SUREilENTS OF TEST SAMPLES 

!,.-.... 

All samples are to be tested and evaluated in accordance with KSC-SPEC-P-0012. 

Prior to installation, each test sample will be measured for an a-curate detennihation of thickness using a template and a .dial gage indicator mounted 

on a flat level table or surface. TG-FLD-22 to determine and rerform the 

measurements. Each test sample will be photographed to record their surface 

condition and texture prior to testing and other conditions detennined as deemed necessary. 

After the launch test, the specimens will be returned to the same area where 

initial measurements wet·e made. The thickness will be remeasured using the 

same techniques to determine the loss of thickness due to general surface erosion. 

Any damage or local spalling or cracking will be noted and recorded. Each 

specimen will be photographed with close-up shots of special conditions. 

The laboratory, TG-FLD-22, will evaluate the relative performance of the 

test sa~ples based upon their findings in accordance with KSC-SPEC-P-0012. 

166 
-. ,.·s· -·--

0/VCE.VT/2 ATEL2 

'vfP11J~EM£1iT 

1TTE.eN .......... 

NASA/TM-2010-216293R 

-·------ ·----·-------- - JOP •• '·-· -

·' 

167 

·--a!-. ---·- . -· 

-=sR 13 FLAM£ 

·.z5£1Z£ciok_ _ --·~- --.--
NASA/TM-2010-216293R 

··~. 

DISTRIBUTION LIST 

SYMBOL/NAME NUMBER OF COPIES 

. PRC-1217/KURTZ 5 

ZK86/VAN DUSEN 1 

PR0:..1214/WALKER 1 

DD~MED~l/OLSEN 1 

---- SI~PR0~14/N. PERRY 1 

SF-ENG/ROBERT MILLER 1 

SF-ENG~J/R • . GILLETT 1 

TG-FLD-2 1 

· TG-FLD-:Z2 · ALL THAT ARE LEFT 

168 
NASA/TM-2010-216293R 

APPENDIX C. MTS-142-82, STS-3 

169 
NASA/TM-2010-216293R 

,/ 

;. NASA 

MATERIALS TESTIN6 SECTION 

MATERIALS ANALY~IS BRANCH 

FLUIDS ANU ANALYSIS OlVlSION 

TG-FL0-22, ROOM 1218, O&C BUILDING 

KENNEDY SPACE CENTER, FLORIDA 32899 

MAY 6, 1982 

MTS-142-82 

SUBJECT: Exposure Test of Refractory Concrete Test Panels to 

Solid Rocket Booster (SRB) Exhaust Uuring STS-3 Launch 

RELATED UOCUHENTATION: MMA-1918-80, July 29, 1981 

MTS-505-81, March 1, 1982 

1.0 FOREWORD 

1.1 On 1 March, 1982, H. G. Olsen, OD-ME0-1, requested 

testing of refractory concrete test panels, the testing 

to comprise thickness measurements, color photographs, 

and visual observations both before and after exposure 

to SRB exhaust during the launch of STS-3. 

1.2 This test is a continuation of the testing reported in 

MMA 1918-80, dated July 29, 1981, and MTS-505-81~ dated 

March 1. 1982. The test panels provid~d by vendors and 

the test fixtures are described in MMA 1918-80, which 

covers the exposure of 6 panels to SRB exhaust during 

STS-1. Exposure of test panels during STS-2 launch 1s 

described in MTS-505-81. 

170 
NASA/TM-2010-216293R 

rn s - 1 4 2 -a 2 2 

2.0 TEST PROCEDURES 

2. 1 TEST PANEL SELECTION AND IDENTIFICATION 

Starting with this test, the positions of the panels in 

the flame trench are designated as A, 8, C, 0, E, and 

F, from east to west in alphabetical order. Of the 

panels exposed during STS-2 launch, panel 6 did not 

survive the launch, and panel 7 was considered too 

badly deteriorated for further testing. The other four 

panels were selected for reexposure during STS-3 

launch. For the other two spaces , panel 1, which had 

been exposed to STS-1 launch, and a previously 

unexposed panel, designated panel 8, were selected. 

The six panels with their exposure positions are as 

follows: 

Exposure 

Position Panel 

East-to-West NO. 

A 8 

B 3 

c 2 

D 

E 4 

F 5 

Designed Concretes, Fondu Fyre 

FSC - 5 

Designed Concretes, Fondu Fyre 

WA-1 w/Wire 

Designed Concretes, Fondu Fyre WA - 1 

Wahl (WRP - 1), contained "fibers" 

(stainless steel wire) 

Wahl (WRP-3), contained carbon 

steel wire 

Harbison-Walker Tufshot, 

contained "fibers" (310 ss) 

2.2 PREEXPOSURE THICKNESS OETERHINATIONS 

Panel 8 was placed on an inspection rack which afforded 

access to both top and bottom surfaces, and thickness 

at each of the predetermined points was measured, using 

171 
NASNTM-20 1 0-216293R 

f.11 S- lt12-82 

the template and caliper employed in previous 

measurements. Photographs of each previously exposed 

panel, made after the last eJ<posure, appear in 

MTS-505-81 and MMA-1918-80. Figure 1 shows panel 8 

before eJ<posure. 

3 

2.3 SRB EXHAUST EXPOSURE 

The panels were then installed in the two refractory 

test frames (DWG 79K09546) at the bottom of the SRB 

flame deflector in the north end of the flame trench at 

LC39A. Each frame holds three panels. The holddown 

c lamps and the areas between panels were coated with 

Dynatherm E-320 ablative coating. The panels were 

installed in the positions indicated in paragraph 2.1, 

and were then exposed to SRB exhaust during STS - 3 

launch. 

2.4 POST-EXPOSURE INSPECTION 

l.4.1 After STS-3 launch, it was found that f i ve of 

the six panels were gone, along with the frames 

and many of the bolts. Two reinforcing grids, 

with fragments of concrete adhering, were found 

hanging on the perimeter fence to the north of 

the pad, Frames, ablative coating, reinforcing 

yr1ds, and some refractory concrete fragments 

were found scattered both Inside and outside the 

perimeter road. 

2.4.2 The surviving panel, Number 1 In position D, was 

remo~ed from the test frame and transported to 

the Inspection rack, where the same template and 

caliper were used to make thickness 

determinations. The depths of representative 

172 
-

NASNTM-2010-216293R 

HTS-142-82 4 

3.0 RESULTS 

spalled areas were determined, using a depth 

gage provided with a dial indicator and a tip 

with a radius of about 1/8 inch. 

3.1 Figure 2 shows panel 1 after launch, and Figures 3 

through 8 show the condition of the test frames and 

sample frames after launch. 

3.2 Panel 1 was dark in color, with scattered light colored 

patches, apparently the result of shallow spalling 

which removed the darker surface layer. The deepest 

pft was 0.270" deep. The average thickness loss during 

this launch was 0.155 inches. Figure 9 shows 

individual thickness measurements before and after both 

STS-1 and STS-3 launches. 

4.0 CONCLUSIONS 

4.1 Only panel 1, Wahl WRP-1, survived the launch. The 

thickness loss and depth of pits were comparable to 

those experienced by this panel during the first 

launch. 

4.2 Two factors wh1ch may hav@ contrfbutQd to th~ holt 

failures whfch resulted fn the loss of sample frames 

and saMples are inadequate cure of the ablative coating 

(Figure 8), and deterioration of the fasteners during 

the first two launches and intervening exposure to 

weather and pad operations. 

173 
NASNTM-2010-216293R 

MTS-142-82 5 

4. 3 Both contractor and government qual fty inspection' wfll 

be requested for future sample installations. This 

should assure complete mixing and adequate application 

of the ablative coating. 

4.4 The sample frames have been redesigned, using larger 

studs which can he replaced when necessary. The grip 

length has been shortened, reducing the bending 

moments. This should provide a longer life and prevent 

s a mp 1 e 1 o s s • 

4.5 Eleven panels w~re received for this investigation. 

Their present. status is as follows: 

P ..a ne 1 

No. 

1 

2* 

3* 

4* 

5* 

Number of 

·Launch 

E X p «!.!1!!:.. e s 

2 (STS-1 

& STS-3) 

3 

3 

3 

3 

-·-------~c rip t ion ______ -·---- Wahl WRP-1. Contained "ffbers" 

(stainless steel wire) 

Designed Concretes, Fondu Fyre 

Wl\-1 

Oesfgned Concretes, Fondu Fyre 

Wl\-1 w/Wire 

Wahl WRP-3. Contained carbon 

steel wire 

Harbison-Walker Tufshot. 

Contained "fibers" (310 ss) 

* Destroyed durfng STS-3 launch • 

..-..._ ** Destroyed durfng STS-2 launch. I 

174 
NASNTM-2010-216293R 

rn" - 1 1 2 - n 2 

Panel 

No. 

tlumbcr of 

Launch 

Exposure~ 

6** 2 

7 

Unttesig- 0 

nated 

B* l 

llnrlesiq- 0 

nated 

Undesig- 0 

nated 

(STS-2) 

(STS-3) 

Harbison-Walker Tufshot 

Wahl WRP-2 (no longer usable) 

Designed Concretes, Fondu Fyre 

W/1-1 (same as Panel 2) 

Designed Concretes, Fondu Fyre 

FSC-5 

Harbison-Walker Tufshot (same as 

Panel 6) 

Harbison-Walker Tufshot with 

"fibers" (310 ss) 

(same as Panel 5) 

* Destroyed during STS-3 launch. 

** Destroyed durfng STS-2 l aunch. 

\ I I /1 INVESTIGI\TOR : (_ ~ ..... ,_le. \; 1.v1 ~~-- -----~- --------------- ------ r.L YDE V. MOYERS 

mno Y£0 ' J!_.__t. j ~1}<,~ {. 4.{_~ -- _, ____ .. __ __ . -~-- __ . C. l. SP~ELD, t.Vr ~F, MTS NASI\ 

175 

6 
NASA/TM-2010-216293R 

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176 
NASA/TM-2010-216293R 

APPENDIX D. MTS-340-82, STS-1, -2, AND 3 

177 
NASA/TM-2010-216293R 

.--

NASA 

MATERIALS TESTING SECTION 

MATERIALS ANALYSIS BRANCH 

FLUIDS AND ANALYSIS DIVISION 

TG-FLD-22, ROOM 1218, O&C BUILDING 

KENNEDY SPACE CENTER, FLORIDA 32899 

JUNE 7, 1982 

MTS-340-82 

SUBJECT: Exposure Tests of Refractory Concrete Test Panels to 

Solid Rocket Booster (SRB) Exhaust During the First 

Three STS Launches: Summary Report 

RELATED DOCUMENTATION: MMA-1918-80, JULY 29, 1981 

MTS-505-82, MARCH 1, 1982 

MTS-142-82, MAY 6, 1982 

1.0 FOREWORD 

1.1 At the request of DD-MED-1, 6 refractory concrete test 

panels have been exposed to SRB exhaust during each of 

the first three STS-launches. 

1.2 Panels were supplied by three vendors. 

1.3 The purpose of the tests was to qualify additional 

suppliers of refractory concrete for use in flame 

trench refurbishment. 

1.4 The three reports listed above in "Related 

Documentation" relate procedures and results of tests 

conducted during each of the three launches. 

1.5 This report summarizes those tests and their results. 

178 

:..f ... _ •• 

.· .... 
NASA/TM-2010-216293R 

2.0 TEST PROCEDURE 

2.1 The thickness of each panel was measured before and 

after each launch exposure at 13 locations determined 

by a template. The maximum depths of pits were also 

determined. 

2 

2.2 The panels were secured in test frames at the bottom of 

the flame deflector in the north end of the flame 

trench of Launch Complex 39A. Figure 1 shows the flame 

trench and test frames. Ablative coating was applied 

to the areas surrounding the panels. Figure 2 shows a 

test frame with panels installed. 

3.0 TEST RESULTS 

3.1 Table 1 summarizes test results for the three launches. 

3.2 Figures 3 through 9 show test panels in the condition 

in which measurements were last made. The graphics in 

the corners of the photographs showing average 

thickness loss and pit depth are the same as those 

included in Table 1. 

4.0 DISCUSSION OF RESULTS 

4.1 For the panels exposed, differences in average 

thickness loss and maximum pit depth were small, with 

two exceptions. Wahl WRP-2, Panel 7, showed much 

greater overall loss than the others. Of the others, 

Harbfson-Walker Tufshot with fibers, Panel 5, showed 

much greater pit depth. 

179 
NASA/TM-2010-216293R 

. ·-

4.2 Harbison-Walker Tufshot, Panel 6, did well in STS-1 

exposure, but was destroyed during STS-2 exposure. 

4.3 Four products have each been subjected to two launch 

exposures with comparably good performances for 

thickness loss and pit depth: 

Designed Concretes Company WA-1 

Designed Concretes Company WA-1 with wire 

Wahl Refractory Products Company WRP-1 

Wahl Refractory Products Company WRP-3 

5.0 CONCLUSIONS. 

3 

5.1 The four products listed in the paragraph above should 

be considered for further testing •. 

5.2 The two Harbfson-Walker ,panels on hand (Tufshot and 

Tufshot with fibers) might be considered for testing. 

INVESTIGATOR :_C _ __::v"--£._~..:.._---'-1~_:___ __________ _ 

C. V. MOYERS 

180 
...... 

00 

...... 

STS-1 

STS-2 

STS-3 

TABLE 1 

DESIGNED CONCRETES CO. WAHL REFRACTORY PRODUCTS COMPANY HARBISON-WALKER REFRACTORIES 

FONDU-FYRE FONDU-FYRE 

WA-1 WA-1 WRP-1 PANEL 1 WRP-2 - PANEL 7 WRP-3 - PANEL 4 TUFSHOT - PANEL 6 TUFSHOT WITH FIBERS 

PANEL 2 WITH WIRE PANEL 5 

PANEL 3 

AVG. LOSS 0. 256 AVG. LOSS 0.325 AVG. LOSS 0.171 NOT EXPOSED AVG. LOSS 0. 235 AVG. LOSS 0.192 AVG. LOSS 0.238 

MAX PIT 0.258 MAX PIT 0.214 MAX PIT 0.222 MAX PIT 0.174 MAX PIT 0.258 MAX PIT 0.346 

CRACKED DURING CRACKED DURING 

INSTALLATION. LAUNCH. NO 

NO ADVERSE ADVERSE EFFECTS 

EFFECTS. 

AVG. LOSS 0.087 AVG. LOSS 0.073 NOT EXPOSED AVG. LOSS 0.257 AVG. LOSS 0.003 PANEL NOT AVG LOSS 0.063 

CUMULATIVE 0.343 CUMULATIVE 0.398 MAX PIT >1.000 CUMULATIVE 0. 238 RECOVERED CUMULATIVE 0.301 

MAX PIT 0.311 MAX PIT 0.201 BADLY ERODED- MAX PIT 0.278 MAX PIT 0.656 

WITHDRAWN FROM SOME OCEP PITS MAY 

TEST. BE ASSOCIATED WITH 

CRACKS. 

PANEL NOT PANEL NOT AVG. LOSS 0.208 NOT EXPOSED PANEL NOT --- PANEL NOT RECOVERED. 

RECOVERED. RECOVERED. CUMULATIVE 0.379 RECOVERED. 

MAX PIT 0.270 

. 0J.· v v ~ \j 'Y' v ----\) I J 

I I 

I 

I i 

u 

2 !3 1 7 4 6 5 

THE l«lRIZONTAL LINES ON THE GRAPHS REPRESENT AVERAGE PANEL THICKNESS AND THE RIGHT PORTION CF EACH GRAPH REPRESENTS MAX PITS 

DEPTH. THE TOP OF· THE GRAPH REPRESENTS THE ORIGINAL PANEl SURFACE. VERTICAL DISTANCES ARE FULL SCALE. 

FOR PANEL 7, THE SOLID LINE REPRESENTS THE VERAGE OF THE POINTS MEASURED. MANY POINTS WERE OUTSIDE THE RANGE (J' THE MEASURING 

EQUIPMENT. THE BROKEN LINE IS ESTIMATED, BUT IS NEARER THE ACTUAL CONDITION. 
....... 

00 

N 

STS· 

STSSTSI 2 

3 

TABLE I 

DESIGNED ClliiCRmS CO. IIAltl REFRACTORY PRODIICTS COIIPAIIY HAR!IISOII-liALl£R REfRACTORIES 

FOIIllU-FYRE F111DU-FYRE 

VA-l VA-l IIRP-1 PANEL I WRP-2 • PANEL 7 WRP-3 

- PANEL 4 TUFSitOT • PAIIEL 6 TUFSHOT II!Tll nBERS 

PAIIEL 2 WITH WIRE PANEL 5 

PANEL 3 

AVG. LDSS 0.256 AVG. LDSS 0.325 AVG. LDSS D.111 IDT EXPOSED AVG. LDSS 0.235 AVG. LDSS 0.192 AVG. LDSS 0. 238 

MAX PIT D.258 MAX PIT 0.214 MAX PIT 0.222 MAX PIT 0.114 W.X PIT D.258 MAX PIT 0.346 

CRACKED OORIIG CRACKED WRING 

IMSTALLATIDN. LAUNCH. NO 

110 ADVERS£ ADVERSE EFFECTS· 

EFfECTS. 

AVG. LDSS 0.087 AVG. LOSS 0.073 NOT EXPOSED AVG. LDSS 0.257 AVG. LDSS 0.003 PANEl lOT AVG LDSS o.ou 

CIIWLATIYE 0.343 CIJIULATIVE 0.398 MAX PIT >1.000 Ctllll.ATIYE 0.238 RECOVERED CIIUATIY£ 0.301 

-

MAX PIT 0.311 MAX PIT 0.201 8ADL T ERODED· MAX PIT 0.278 W.X .PIT 0.656 

WITHDRAWN FROM SOME I££P PITS W.Y 

TEST. BE ASSDC lA TED IIITH CRACKS. 

PANEL lilT PAIIEL lilT AVG. LDSS 0.208 NOT EXPOSED PAVEL lOT -- PANEL lOT RECOVERED. 

RECOVERED. RECOVERED. CIMULATIVE 0.379 RECOVERED. 

MAX PIT 0.270 

~ "\::1 v \1 -~ ~ 

----~ I . . \ ' I I : \ I 

I I 

u 

2 3 I 7 4 6 5 

-- ·-- --

- --- --- ---- --- . --- --- -· ··-·· 

- --··. ·- ·-- -- ··-

THE lllRIZIJITAL LINES ON THE GRAPHS REPRESENT AVERAGE P.IIIEL TliiCKNESS Nlll 1HE RIGHT PORTIOII IF EACH GRAPH REPRESENTS MAX "PITS 

D£PTll. TilE TDP OF THE GRAPH REPRESENTS THE ORIGINAL PANEL SURFAC£. VERTICAL DISTANCES Alit: FULL SCALE. 

FOR PANEL 7, THE SOLID LINE ·REPRESENTS TIE VEAAGE IF TIE PDIITS MEASURED. MAllY POIITS WERE OUTS liE TIE RANGE IF TIE MEASURIIG ., 

EQUIPMENT. TilE BROKEN LINE IS ESTIMATED, BUT IS NEARER TilE ACTUAL COIID!TIOII. 
NASA/TM -201 0-216293R 

5 

FIGURE 1 

FLAM E TRENCH, SHOWING TEST FRAMES AT BOTTOM OF FLAME DEFLECTOR 

183 
NASA/TM-2010-216293R 

6 

FIGURE 2 

TEST FRAME WITH TEST PANELS INSTALLED 

184 
-- --- -·------------------------

NASNTM-2010-216293R 

7 

2 

FIGURE 3 

DESIGNED CONCRETES WA-1 AFTER STS-2 

185 
NASA/TM-2010-216293R 

8 

FIGURE 4 

DESIGNED CONCRETES WA-1 W/W AFTER STS-2 

186 
- - -------- - --- - ---

NASA/TM-2010-216293R 

9 

FIGURE 5 

WAHL WRP-1 AFTER STS-3 

187 
NASNTM-2010-216293R 

10 

FIGURE 6 

WAHL WRP-2 AFTER STS-2 

188 
NASA/TM-2010-216293R 

11 

FIGURE 7 

WAHL WRP-3 AFTER STS-2 

189 
NASNTM-2010-216293R 

FIGURE 8 

HARBISON-WALKER TUFSHOT AFTER STS-1 

190 

12 
NASA/TM-2010-216293R 

13 

FIGURE 9 

HARBISON-WALKER TUFSHOT W/W/ AFTER STS-2 

191 
NASNTM-2010-216293R 

.---

__,· 

DISTRIBUTION LIST 

SYMBOL/NAME NUMBER OF COPIES 

PRC-1217/KURTZ 5 

ZK86/VAN DUSEN 

PR0-1214/WALKER. 

DD-MED-1/0LSEN 5 

SI-PR0-14/N, PERRY 

-.../ 

SF-ENG/ROBERT MILLER 

SF-ENG-3/R. GILLETT 

TG-FLD-2 

TG-FLD-22 ALL THAT ARE LEFT 

-. 

-------- ------------· 

192 
NASA/TM-2010-216293R 

APPENDIX E. MTS-425-82, STS-4 

193 
NASA/TM-2010-216293.R 

...... 

NASA 

MATERIALS TESTING SECTION 

MATERIALS ANALYSIS BRANCH 

FLUIDS AND ANALYSIS DIVISION 

TG-FLD-22, ROOM 1218, O&C BUILDING 

KENNEDY SPACE CENTER, FLORIDA 32899 

AUGUST 27, 1982 

SUBJECT: Exposure Test of Refractory Concrete Test Panels During 

STS-4 Launch 

RELATED DOCUMENTATION: MTS-340-82 Exposure tests of Refractory 

Concrete Test Panels to Solid Rocket 

Booster {SRB) Exhaust During the First 

Three STS Launches 

1. 0 FOREWORD 

This test was a continuation of the test program described 

in MTS-340-82. ·Six test panels were prepared by the 

Development Testing Branch, using material from two 

manufacturers. 

2.0 TEST PROCEDURE AND RESULTS 

The test procedure was similar to that followed in 

previously reported tests of this kind. Results are as 

fo 11 ow s: 

·-------·-··--·-----·--·----------·------ --------------------- ----------·- ----· ---------.------------

194 
.. -·-.. 

NASA/TM -201 0-216293R 

MTS-425-82 2 

% OF 

LOCATIONS 

WITH 

APPARENT AVERAGE 

BRAND/ PANEL EXPOSURE MAXIMUM THICKNESS THICKNESS 

.MFGR NUMBER POSITION PIT DEPTH (INS) INCREASE* LOSS (INS) 

FONOU 11 B 0.325 0 0.048 

FYRE 

13 D 0.700 0 0.178 

(.1. 00 AT CRACK) 

15 F 0.227 15 0.047 

WAHL 10 A 0.142 62 -0.011 

(INCREASE 

IN 

THICKNESS) 

12 c 0.361 15 o. 077 

14 E 0.180 15 0.022 

* Thickness was measured at 13 locations on each panel. 

Cracks were noted in several panels, but did not appear to 

be significant except in panel 13, in which the cracks had 

been widened and deepened by erosion. 

3.0 DISCUSSION 

3.1 The condition of panel 13 compared with that of the 

other two of the same composition exemplifies the 

uneven exhaust exposure at different locations across 

the flame trench which has been observed in previous 

launches. 

-·--·------·-·- ..... ······-·---. ··--···----······ ··-·--·-·--·- --------·- ----·-·--------------·-----

195 
NASA/TM-2010-216293R 

.. MTS-425-82 3 

3.2 Apparent increases in thickness were measured in a 

number of locations, especially on the Wahl panels. As 

in some previous occurrences, they are attributed to 

the accretion of exhaust products. 

IN VE s TI GATOR :___:;C::_. _h+-ck-=-::..__V__:_m_. '+l-~-=---------- CLYDE V. MOY.ERS 

.- ----- -- ---------------- ~---

196 
NASA/TM -201 0-216293R 

TASK REQUEST 

I. DATE SUBMITTED 2. DESIRED COMPLETION DATE 8. AUTHORIZING DOCUMENT 

SAMPLE DESCRIPTION: 

15, SYSTEM REMO YEO FROM/OR USED IN: 

4, ANALYSIS REQUESTED! 

8, REQUESTER: 

~ASA APPROVAL: 

I'I.SC FORM 22-81 (REV, 1/70) 

197 
NASA/TM-2010-216293R 

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198 
- --------------- --- -------- ---- ------- 

NASA/TM-2010-216293R 

APPENDIX F. MTB-503-83, STS-5, -6, AND -7 

199 
NASA/TM-2010-216293R 

NASA· 

ENGINEERING DEVELOPMENT 

GROUND SUPPORT OFfiCE 

MATERIALS ANALYSIS OFFICE 

MATERIALS TESTING BRANCH 

DE-MA0-2, .ROOM 1218, O&C BUILDING, PHONE 867-4614 

KENNEDY SPACE CENTER, FLORIDA 32899 

SEPTEMBER 1, 1983 

MTB-503-83 

SUBJECT: Exposu~e Tests of Refractory Concrete Test panels to 

Solid Rocket Booster{SRB) Exhaust During the First 

Seven STS Lanches: Summary Report 

RELATED DOCUMENTATION: MMA-1918-80, JULY 29, 1981 

MTS-505-82,,MARCH 1, 1982 

MTS-142-82, MAY 6, 1982 

MTS-340-82, JUNE 7, 1982 

MTS-425-82, AUGUST 27, 1982 

1.0 FOREWORD 

1.1 At· the request of DD-MED-1, refractory concrete test 

panels have been exposed to SRB exhaust during each of 

the first seven STS-launches. 

1.2 Panels were supplied by three vendors, except for six 

panels which were fabricated at KSC from 

vendor-supplied materials. 

1.3 The purpose.of the test was to qualify additional 

suppliers of refractory concrete for use in flame 

trench refurbishment. 

200 
NASA/TM-2010-216293R 

MTB-503-83 2 

1.4 The five reports listed above in "Related 

Documentation" relate procedures and results of tests 

conducted during the first four launches. 

1.5 This report summarizes the results of those tests, as 

well as the hitherto unreported tests during STS-5, 

STS-6 and STS-7 launches. 

2.0 TEST PROCEDURE 

2.1 The panel thickness was measured before and after each 

launch exposure at 13 locations on each panel. A 

template was used to identify the locations to be 

measured. The maximum depths of pits were also 

determined. 

2.2 For launch exposure the panels were secured in test 

frames at the bottom of the flame deflector in the 

north end of the flame trench of Launch Complex 39A. 

Panel positions were designated A through F in 

alphabetical order from east to west. Ablative coating 

was applied to the areas surrounding the panels to 

protect the fixtures used to secure the samples in 

place. 

3.0 RESULTS 

3.1 Table 1 summarizes test results for the seven 

launches. 

results. 

Table 2 is a more detailed tabulation of 

3.2 Figures 1 through 3 show graphically the effects of 

launch exposure on individual panels. 

201 
NASNTM-2010-216293R 

MTB-503-83 3 

4.0 DISCUSSION OF RESULTS 

The material now in use in the flame trench is Designed 

Concretes' Fondu Fyre WA-1. The results for this material 

shown in Figure 1 represent five panels with a total of 

seven launch exposures. Three of these panels were 

fabricated at KSC. Three other materials show lower average 

thickness losses per launch than WA-1: 

4.1 Harbison Walker Refractories 17-67 

The average loss is low, but represents only two 

exposures of one panel prepared by the vendor. 

4.2 Wahl Refractory Products WRP-3 (Wi~h Wire) 

The average loss is low, and represents eleven 

exposures of six panels, three of which were prepared 

at KSC. 

4.3 Wahl Refractory Products WRP-3 without wire 

The average loss is low, but represents only two 

exposures of one panel prepared by the vendor. 

4.4 In a memorandom to M.G. Olsen, DD-MED-1, dated 

August 22; 1983 Gary Kurtz, PRC 1211, has reviewed the " test program so far, and has discussed the problems in 

test exposure and interpretation of results. In the 

evaluation of results he suggests the following 

criteria: 

"a. eliminate specifit materials from contention 

that crack or form deep fissures. Exceptions to 

this rule would be panels that have knowingly 

been mishandled or otherwise, improperly 

installed. This would be consistent with the 

requirements as stated in KSC-SPEC-P-0012, 

section 3.5. 

202 
---- ----

NASA/TM-2010-216293R 

MTB-503-83 4 

b. eliminate maximum spalltng depth from the list 

of SPEC requirements. It appears to be basic 

property of the refractories that they all 

spall. 

c. use a minimum of two panels of a specific 

material for a minimum of three launch 

exposures. One of the panels should be prepared 

by NASA, KSC personnel or its contractor. 

d. use Fondu Fyre WA-1 average loss as a minimum 

standard of acceptability. This may seem 

somewhat arbitrary, but then so was the 1/8" 

loss quoted in the SPEC. Attempts to arrfve at 

an acceptable figure for loss (other than zero) 

all seem subjective and without any real 

meaning. The only practical vi ewpoint with the 

information available at this time is that Fondu 

Fyre WA-1 did well but it would be beneficial to 

find better materials. Cost effectiveness has 

not been considered so far, however, material 

cost, labor, scheduling, and availability will 

have to be factored into the acceptability 

equation at some point." 

Using these criteria, he says, "the only material that 

would qualify for usage, other than Fondu Fyre WA-1, 

would be Wahl WRP-3 (w/w)." But he cautions that 

compatibility of the two materials and their 

performance as patches in repair work have not been 

investigated. Both Wahl WRP-3 without wire and 

Harbison Walker 17-67 would require additional 

successful launch exposures, including exposures of 

KSC-fabricated panels, before being qualified. 

203 
NASNTM-2010-216293R 

MTB-503-83 5 

Mr. Kurtz suggests future work to study cost 

effectiveness, compatibility, application methods, new 

materials, and variables in material preparation. 

5.0 RECOMMENDATIONS 

Mr. Kurtz's analysis seems reasonable and comprehensive. 

Two alternate courses are suggested: 

5.1 Exposure tests have now been conducted during seven 

launches over a period of more than 2 years. If the 

need for an additional qualified vendor is urgent 

enough, then Wahl WRP-3 could be accepted on the basis 

of the criteria outlined in paragraph 4.4. In this 

case the assumption would be made that any problems. 

which may arise in repairs and compatibility will be 

solved by appropriate application technique. Exposure 

tests of Harbison Walker 17-67 and Wahl WRP-3 without 

wire with additional panels fabricated at KSC would 

continue. Mr. Kurtz's suggestions should be considered 

in planning future work. 

5.2 Alternatively, acceptance of a new product could be 

postponed pending repair and compatibility testing. In 

this case, a test plan should be made immediately, 

·204 
MTB-503-83 

NASNTM-2010-216293R 

6 

using test panels, flame deflector patches, or both. 

Exposure tests of Harbison Walker 17-67 and Wahl WRP-3 

without wire, using panels fabricated at KSC, would 

continue, and future work would be planned as in the 

first alterna t ive. 

INVESTIGATOR : c ~ ~~U} C. V. MOYERS 

MTB, NASA 

205 
NASA/TM-2010-216293R 

MTB-503-83 

TABLE 1 

SUMMARY-LAUNCH EXPOSURES 

AVERAGE 

NUMBER THICKNESS 

OF LAUNCH LOSS PER 

VENDOR EXPOSURES LAUNCH, INS. 

DESIGNED FONDU FYRE WA-1 7 

CONCRETES FONDU FYRE WA-1 WITH WIRE 2 

.HARBISON TUFSHOT 1 

WALKER TUFSHOT.WITH FIBER 2 

REFRACTORIES LI 1 

17-67 2 

WAHL WRP-1 2 

REFRACTORY WRP-2 1 

PRODUCTS WRP-3 (WITH WIRE) 11 

.WRP-3 WITHOUT WIRE 2 

206' 

0.121 

0.199 

0.192 

. 0.151 

0.156 

0.068 

0.190 

0.257 

0.055 

0.057 

. 7 

AVERAGE 

MAXIMUM 

PIT DEPTH, 

INS. 

0.347 

0.208 

0.258 

0.501 

0.475 

0.368 

0.246 

>1.000 

0.231 

0.468 
NASA/TM-2010-216293R 

MTB-503-83 8 

TABLE 2 

LAUNCH EXPOSURES 

AVERAGE MAXI t-UM 

PANEL THICKNESS PIT DEPTH 

\I ENDOR DESIGNATION NUMBER LAUNCH POSITION LOSS, IN. IN. 

UESIGNED FONDU FYRE WA-1 2 STS-1 B 0.256 0.258 

CONCRETES STS-2 E .087 .311 

18 STS-6 A .041 .159 

STS-7 c .189 .450 

11* STS-4 B .048 .325 

STS-5 E ( .038**) 

13* STS-4 D .178 .700 

15* STS-4 F .047 .227 

AVERAGE PER LAUNCH - - - - - - - - 0.121 0.347 

FONDU FYRE WA-1 3 STS-1 c 0.325 0.215 

WITH WIRE STS-2 F .073 .201 

AVERAGE PER LAUNCH - - - - - 0.199 0.208 

HARBISON TUFSHOT 6 STS-1 F 0.192 0.258 ~IALKER TUFSHOT WITH FIBER 

5 STS-1 E 0.238 0.346 

STS-2 B .063 .656 

AVERAGE PER LAUNCH - - - ----- 0.151 0.501 

LI 21 STS 6 c 0.156 0.475 

17-67 22 STS-6 E 0.037 0.375 

STS-7 E .098 .360 

AVERAGE PER LAUNCH - - - - - - 0.068 0.368 

*FABRICATED AT KSC *ABOUT 15% OF PANEL SURFACE MISSING DOWN TO REINFORCEMENT. VALUE REPORTED MEASURED OVER EMAINING 85% - NOT INCLUDED IN AVERAGE PER LAUNCH. 

207 
NASNTM-2010-216293R 

MTB-503-83 9 

TABLE 2 (CONT'D) 

LAUNCH EXPOSURES 

AVERAGE MAXIMUM ~ANEL THICKNESS PIT DEPTH VENDOR DESIGNATION NUMBER LAUNCH POSITION LOSS, IN. IN. 

WAHL WRP-1 1 STS-1 A 0.171 0.222 REFRACTORY STS-3 D 0.208 0.270 PRODUCTS 

AVERAGE PER LAUNCH - ------- 0.190 0.246 

WRP-2 7 STS-2 D 0.257 >1.000 

WRP-3 4 STS-1 D 0.235 0.174 

STS-2 A 0.003 0.278 

10* STS-4 A -.011 0.142 

STS-5 D 0.072 0.171 

12* STS-4 c 0.077 0.361 

14* STS-4 E 0.022 0.180 

STS-5 B 0.019 0.160 

STS-6 F 0.045 0.140 

STS-7 B 0.044 0.180 

20 STS-6 D 0.049 0.280 

STS-7 F 0.046 0.480 

AVERAGE PER LAUNCH - - ------ 0.055 0.231 

WRP-3 WITHOUT WIRE 19 STS-6 B 0.019 0.600 

STS-7 D 0.095 0.335 

AVERAGE PER LAUNCH - - - - - - - - 0.057 0.468 

*FABRICATED AT KSC 

208 
MTB-503-83 

FON{lU FYRE 

WA-1 · PANEL 2 · 

.. .. . , . 

FONDU FYRE 

WA-1 

PANEL 18 

. FONQU _FYRE 

WA.:. l . ·PANEL is*. 

FIGURE 1 

FONDU FYRE 

WA-i 

PANELll* ,** 

. FONDU FYRE 

WA-1 WJTH WIRE 

PANEL 3 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. 

DESIGNED CONCRETES COMPANY 

NASA/TM-2010-216293R 

tO 

FONDU FYRE 

WA-1 ·· PANEL 13• 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND MAXIMUM PIT DEPTHS FOR INDIVIDUAL 

LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED ON PANEL 2: 

1. ORIGINAL PANEL SURFACE 

2. AVERAGE SURFACE LEVEL AFTER FIRST LAUNCH EXPOSURE. THE VERTICAL DISTANCE 

FROM 1 TO 2 IS A FULL SCALE REPRESENTATION OF THICKNESS LOSS DURING THIS 

EXPOSURE. 

3. AVERAGE SURFACE LEVEL AFTER SECOND LAUNCH EXPOSURE. THE VERTICAL 

DISTANCE BETWEEN 2 AND 3 IS A FULL SCALE REPRESENTATION OF THICKNESS LOSS 

DURING THE SECOND EXPOSURE. 

4&5 MAXIMUM PIT DEPTHS MEASURED AFTER FIRST AND SECOND EXPOSURES. AS IN "1" 

AND "2", DEPTHS ARE SHOWN IN FULL SCALE • 

.>ANEL FABRICATED AT KSC. ~*SECOND EXPOSURE OF PANEL 11 NOT INCLUDED IN AVERAGES IN TABLE 1 BECAUSE OF POSSIBLE 

ADVENTITIOUS MECHANICAL DAMAGE AND DIFFICULTY OF EVALUATION. 

209 
NASNTM-2010-216293R 

MTB-503-83 

TUFSHOT 

PANEL 6 

HARBISON WALKER REFRACTORIES 

TUFSHOT 

WITH FIBERS 

PANEL 5 

FIGURE 2 

LI 

~ANEL 21 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS~ 

HARBISON WALKER REFRACTORIES. 

1\ 

17-67 

PANEL 22 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND MAXIMUM PIT DEPTHS 

FOR INDIVIDUAL LAUNCH EXPOSURES. TYPICAL FEATURE~ ARE INDICATED 

AND EXPLAINED FOR PANEL 2, FIGURE li Q.V. 

210 

'1 ..... j 

I 

' 

.i 

. ' 

, . I 

•t 
-·-·---·------- -------------

NASA/TM-2010-216293R 

MTB-503-83 12 

I . 

' WAHL REFRACTORY PRODUCT~---

WRP-1 . PANEL 1 

WRP-3 . PANEL 12* 

WRP-2 

PANEL 7 

WRP-3 . PANEL 14* 

FIGURE 3 

WRP-3 

PANEL 4 

WRP-3 

PANEL 20 

.WRP-3 

PANEL 10* 

WRP-3 

PANEL19 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. WAHL 

REFRACTORY PRODUCTS. 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES ANO MAXIMUM PIT DEPTHS 

FOR INDIVIDUAL LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED 

AND EXPLAINED FOR PANEL 2, FIGURE 1, Q.V. 

*PANEL FABRICATED AT KSC. 

211 
NASNTM-2010-216293R 

MTB-503-83 

DISTRIBUTION LIST 

OFFICE SYMBOL/NAME 

D D- ME D- 1 I M. 0 L SEN 

SF-ENG/ROBERT MILLER 

SF-SEC-3/RONALD GILLETT 

SI-PR0-14 

ZK86/VAN DUSEN 

PRC~1372/J. WALKER 

PRC-1211/GARY KURTZ 

DE-MAO/C.W.HOPPESCH 

D D- ME D- 3 3 I J • F • M C I N E R N Y 

SL-ENG-1/WAYNE PARRIS 

NASA, GODDARD SPACE FLIGHT CENTER 

GREENBELT, MARYLAND 20771 

ATTN: MR. RICHARD MARRIOTT- CODE 313 

DE-MA0-2 

NO. OF COPIES 

5 

1 

1 

1 

1 

1 

1 

1 

1 

1 

ALL THAT ARE 

LEFT 

PLEASE NOTIFY THIS OFFICE (867-4614/DE-MA0-2) IF THERE ARE ANY 

CHANGES IN NAME, ADDRESS, OFFICE SYMBOl, ETC. -THANK YOU! 

212 
NASA/TM-2010-216293R 

~--------------------------------------------------------------------------------~ I 

DATE SUBMITTED: 

1121193 

TASK REQUEST 

DES I RED COMP DATE: 

214193 

AUTHORIZING DOC: 

1 SAMPLE DESCRIPTION: 

I 

.I 

-

Six (6) samples of material chipped from refractory concrete test 

specimens: 

Sam2le No. Identification 

1. Fondu Fyre TIS 1. MS 3 

2. Mitec 1111A05071. TIS 2. MS 1. 

3 . Mitec AIO. TIS 3 . MS 5. 

4. Mitec TIS 4. MS 6. 

5 . Fondu Fyre TIS 5, MS 4. 

6. Mitec 11215071, TIS 6, MS 2. 

--· 

SYSTEM: Evaluation of Refractory Concrete, Flame Deflector. LC-39A 

ANALYSIS REQUESTED: 

Preliminary Petroqraph i c Analvsis: 

Perform Semi-quanti tative analysis of particulate. 

REMARKS: 

REQUESTOR: I PHONE: I COMPANY: I MAIL CODEr 

P. J. Welch 867-1403 NASA DM-MSL-24 

FOR LAB USE ONLY 

INVESTIGATOR I TASK NO. I SAMPLE NO. 

213 
NASA/TM-2010-216293R 

This page intentionally left blank. 

214 
NASA/TM-2010-216293R 

APPENDIX G. MTB-250-84, STS-8 AND -9 

215 
NASNTM-2010-216293R 

NASA 

ENGINEERING DEVELOPMENT DIRECTORATE 

GROUND SUPPORT OFFICE 

MATERIALS ANALYSIS OFFICE 

MAfERIALS·.TESTING BRANCH 

DE~MA0-2, ROOM 1218, O&C BUILDING, PHONE 867-4614 

kENNEDY SPACE CENTER, FLORIDA 32899 

MAY 22, 1984 

SUBJECT: Exposure Tests of Refractory Concrete Test panels to 

Solid Roc,et Booster(SRB) Exha~st Durin~ the First 

Nine STS launches: Summary Report 

RELATED DOCUMENTATION: MMA-1918-80, JULY 29,'198..;-· 

'l -;;> HTS-505-82, MARCH 1, 1982 ~ 

1. 0 FOREWORD 

MTS-142-82, MAY 6, 198211 

MTS-340-82, JUNE 7, 1982/ 

MTS-425-82, AUGUST 27, 1982.1 

MTB-503-83, SEPTEMBER 1, 19.3 ~ 

1.1 At the request of DD-HED-1, refractory concrete test 

pa~e1s have been exposed to SRB exhaust du~1ng each ~f 

the first nine STS-launches. 

1.2 Panelf were supplied 'Y four vendor•, except fof i1x 

panels which were fabricated at KSC from 

vendor-supplfed materials. 

1,3 The purpose of the test was to qualify additional 

suppliers of refractory· concrete for use 1n flame 

trench refurbishment. 

216 

t. 
NASA/TM-2010-216293R 

HTB-250 - R4 2 

1.4 The six repor~s 11st~d above in "Related Documentation" 

relate procedures and 'results of tests conducted during 

the first seven launches. 

1.5 Thfs report summarizes the results of those tests, as 

well as the hitherto unreported tests during STS-R and 

S TS- 9 launches, 

2.0 TEST PROCEDURE 

2.1 The panel thickness was measured before and after each 

Iaurich exposure at 13 locations on eAch panel. A 

template was used to identify the locations to be 

measured, The maximum depths of pfts were also 

determined, 

2.2 For launch exposure the panels were secured fn test 

frames at the bottom of the flame deflector fn the 

north end of the flame trench of launch Complex 39A. 

Panel posftfons were designated A through F fn 

alphabetical order from east to west. Ablative coating 

was applied to the areas surrounding the panels to 

protect the fixtures used to secure the samples in 

place. 

3,0 RESULTS 

3.1 Table 1 sum~arizes test results for the nine launches. 

Table 2 is a more detailed tabulation of results, 

3.2 Figures 1 through 4 show graphically the effects of 

launch exposure on individual panels. 

217 
NASA/TM-20 1 0-216293R 

HTB-250-1\4 3 

4.0 OJSCUSSJON OF RESULTS 

The m~ter1al now 1n use 1n the flame trench 1s Designed 

Concretes' Fondu Fyre WA-1. The results for this material 

shown in Figure 1 represent five panels with a total of 

seven launch exposures. Three of these panels were 

fabri~ated at kSC. Four other materials show lower average 

thickness losses per launch than WA-1: 

4.1 Harbison Walker Refractories 17-67 

The average loss is low, but represents only three 

exposures of one panel prepared by the vendor. 

4.2 Wahl Refractory Products WRP-3 (With Wire) 

The average loss is low, and represents thirteen 

exposures of six panels, three of which were prepared 

at kSC. 

4.3 Wahl Refractory Products WRP-3 without wire 

The average loss 1s low, but represents only two 

exposures of one panel prepared by the vendor. 

4.4 Sauere1sen Cements Company No. 75 

I 

The average loss 1s low, but represents only two panels 

exposed to STS-9. These panels were prepared by the 

vendor, and contain chopped wire, but no Gr1dstee1 

reinforcement. One panel was covered with a network of 

fine cracks after the launch, and the other had a few 

cracks, but both felt solid and withstood moving by 

truck and forklift and being placed on inspection racks 

where they were supported on two opposite edges without 

further visible damage. 

218 
NASA/TM-2010-216293R 

HTB-250-R4 

5.0 PROJECTED TESTING 

Four panels are scheduled for exposure during the next 

launch, 41-D. The edge positions, A and F, which receive 

less exhaust exposure, will not by used. 

5.1 _lwo previously exposed WA-1 panels wfll have a strip 

sawn out and w111 be repaired, using WA-1 for one 

repair and WRP 3 for the other. 

5.2 An unexposed WA-1 panel w111 be exposed to provide the 

basfs for a later repair test. 

5.3 A Sauerefsen No. 75 panel wfll be prepared at KSC, 

using wfre and Grfdsteel refnforce~ent, and w111 be 

exposed. 

J NVEST J GATOR: c. V/Y1,;.e--n 

C. Y. MOYERS 

APPROYED:~l?.~,~~~~~~~~~~---------------------­ C. l. 

219 
NASA/TM-2010-216293R 

-

---

r 

MTB-250-84 

VENDOR 

DESIGNED 

CONCRETES 

'-

HARBISON 

WALKER 

REFRACTORIES 

WAHL 

REFRACTORY 

PRODUCTS 

SAUEREISEN 

COHENTS CO. 

TABLE 1 

SUMMARY-LAUNCH EXPOSURES 

FONDU. FYRE WA-1 

FONDU FYRE WA·1 WITH WIRE 

FONDU FYRE FSC-5 

FONDU FYRE HT-1 

TUFSHOT 

TUFSHOT WITH FIBER 

Ll 

17-67 

WRP·l 

WRP·2 

WRP-3 (WITH WIRE) 

WRP-3 WITHOUT WIRE 

NO~ 75 

NUMBER 

OF LAUNCH 

EXPOSURES 

1' 

2 

1 

1 

1 

2 

1 

3 

2 

1 

13 

2 

2 

220 

AVERAGE 

THICKNESS 

LOSS PER 

LAUNCH. INS. 

0.121 

0.199 

0.103 

0.172 

0.192 

0.151 

0.156 

0.080 

0.190 

o.257 

0.053 

0.057 

0.069 

5 .. 

AVERAGE 

MAXIMUM 

PIT DEPTH, 

INS. 

0.347 

0.208 

0.270 

0,280 

0,258 

0.501 

0.475 

0.455 

' 0.246 

>1•000 

0.246 

0.468 

0.370 
NASNTM-2010-216293R 

HTB-250 - 84 6 

TABLE 2 

LAUNCH EXPOSURES 

AVERAGE MAXIMUM 

PANEL THICKNESS PIT DEPTH 

VENI>OR DESIGNATION NUMBER LAUNCH POSITION LOSS, IN. IN. 

DES .IGNED FONDU FYRE WA-1 2 STS-1 B 0.256 0.258 

CONCRETES STS-2 E 0.087 o. 311 

18 STS-6 A 0.041 0.159 

STS-7 c 0.189 0.450 

11* STS-4 B 0.048 0.325 

STS-5 E (0.038**) 

13* STS-4 D 0.178 0.700 

15* STS-4 F 0.047 ·0.227 

AVERAGE PER LAUNCH • . -.. - - 0.121 0.347 

FONDU FYRE WA-1 3 STS-1 c 0.325 0.215 

WITH WIRE STS-Z F 0.073 0.201 

AVERAGE PER LAUNCH - -.. -- 0.199 0,208 

FONDU FYRE FSC-5 23 STS-9 c 0.103 0.270 

FONDU FYRE HT-1 24 STS-9 D 0.172 0.280 

HARIBJSOH TUFSHOT 6 STS-1 F 0.192 0,25R 

WALKER TUFSHOT WITH FIBER 

5 STS•l E O.l36 0.346 

STS-2 B 0.063 0.656 

AVERAGE PER LAUNCH - - - - - 0.151 0.501 

LJ 21 STS 6 c 0.156 0.475 

17-67 22 STS-6 E 0.037 0.375 

STS-7 E 0.098 0,360 

STS-8 c 0.105 0.630 

AVERAGE PER LAUNCH - ---- 0.068 0.368 

~BRICATED AT ICSC 

**ABOUT 151 OF PANEL SURFACE MISSING DOWN TO REINFORCEMENT. VALUE REPORTED MEASURED OYER 

REMAINING 851 - HOT INCLUDED IN AVERAGE PER LAUNCH. 

221 
NASA/TM-2010-216293R 

MT8-250-84 7 

TABLE 2 (CONT'D) 

LAUNCH EXPOSURES 

AVERAGE MAXIMUM 

.DESJGNATii:JN ~ 

PANEL THICKNESS PIT DEPTH 

~ LAUNCH POSITION LOSS, IN. IN, 

WAHL WRP-1 1 . STS-1 A 0,171 0.222 REFRACTORY STS-3 D 0.208 0.270 PRODUCTS 

AVERAGE PER LAUNCH - ------- 0.190 0.246 

WRP-2 7 STS-2 D .. 0.257 >1.000 

WRP-3 4 STS-1 0 0,235 0,174 

STS-2 A 0.003 0.278 

,- ,.) 

10* STS-4 A .:.on 0.142 

STS-5 0 0,072 0.171 

'12* 'STS-4 c 0.077 0,361 

14• STS-4 E 0.022 0,180 STS-5 8 0.019 0.160 

STS-6 F 0,045 0.140 

STS-7 8 0.044 0.180 

STS-8 8 0,070 0.130 

20 STS-6 D 0.049 0.280 

STS-7 F 0.046 0.4RO 

STS-8 E 0,017 0,530 

AVERAGE PER LAUNCH .: - - - • • 0.053 0.246 

WRP-3 WITHOUT' WIRE 19 STS-6 B 0.019 0.600 

STS-7 D 0,095 0,335 

AVERAGE PER LAUNCH •-- ··•- ~- 0.057 0,468 

SAUEREISEN N0,'75 '25 STS-9 B 0,103 0.410 

.(~- 'ENTS CO• 26 STS-9 E 0.035 0,J30 

' 'AVERAGE PER LAUNCH • ~ - ·• • • - • Q.069 o·:370 

'*FAB~ICATED AT kS~· 

222 
NASNTM-2010-216293R 

MTB-250-84 

FOtmu· FYRE 

WA-1 

PAII[L 2 

FONQU FYRE 

WA-1 ·PANEL 15*. 

DESIGNED CONCRETES CO. 

f{)NDU FYRE 

WA-1 · PANEL 1B 

. FONDU FYRE 

WA-1 WITH WIRE 

PANEL 3 

FIGURE 1 

1V 

FONOU FYRE 

WA-1 

PANELll*•** 

J 

. '\._/ 

FONDU FYRE 

FSC-5 

PANEL 23 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. 

DESIGNED CONCRETES COMPANY 

8 

FONDU FYRE 

WA-1 

PANEL 13* 

FONDU FYRE 

HT-1 

PANEL 24 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND MAXIMUM PIT DEPTHS FOR INDIVIDUAL 

LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED ON PANEL 2: 

1. ORIGINAL PANEL SURFACE 

2. AVERAGE SURFACE LEVEL AFTER FIRST LAUNCH EXPOSURE. THE VERTICAL DISTANCE 

FROM 1 TO 2 IS A FULL SCALE REPRESENTATION OF THICKNESS LOSS DURING THIS 

EXPOSURE. 

3. AVERAGE SURFACE LEVEL AFTER SECOND LAUNCH EXPOSURE. THE VERTICAL 

DISTANCE BETWEEN 2 AND 3 IS A FULL SCALE REPRESENTATION OF THICKNESS LOSS 

DURING THE SECOND EXPOSURE. 

4&5 MAXIMUM PIT DEPTHS MEASURED AFTER FIRST AND SECOND EXPOSURES. AS IN •1• ~ AND •z•, DEPTHS ARE SHOWN IN FULL SCALE. 

I' *PANEL FABRICATED AT kSC. 

•*SECOND EXPOSURE OF PANEL 11 HOT INCLUDED IN AVERAGES IN TABLE 1 BECAUSE OF POSSIBLE 

1'DYENTI TIOUS MECHANICAL DAMAGE AND DIFF ICUL TV Of EVALUATION. 

223 
NASNTM-2010-216293R 

HTB-250-84 

TUFSHOT 

PANEL 6 

HARBISON WALkER REf'RACTORIES 

TUF'SHOT 

WlTH FIBERS 

PANEL-S 

FIGURE 2 

LJ 

PAfiEL 21 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. 

HARBISON WALkER REFRACTORIES. 

' 

17-67 

PANEL 22 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND MAXIMUM PIT DEPTHS 

FOR INDIVIDUAL LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED 

AND EXPLAINED FOR PANEL 2, FIGURE 1, Q.V. 

224 
-

I 

NASA/TM-2010-216293R 

MTB-250- 84 

HRP-1 

PANEL 1 

WRP-3 . PANEL -12* 

WAHL REFRACTORY PRODUCTS 

WRP-2 

PANEL 7 

WRP-3 

PANEL 14* 

FIGURE 3 

WRP-3 

PANEL 4 

WRP-3 

.PANEL 20 

.WRP-3 

PANEL 10* 

WRP-3 

PANEL19 · 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. WAHL 

REFRACTORY PRODUCTS. 

.10 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND MAXIMUM PIT DEPTHS 

FOR INDIVIDUAL LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED 

(,.- AND EXPLAINED FOR PANEL 2. FIGURE 1. Q.Y. 

*PANEL FABRICATED AT KSC. 

225 
NASA/TM-2010-216293R 

MTB-250-84 

SAUREJSEN CEMENTS CO . 

v 

NO. 75 

PANEL 25 

FIGURE 4 

NO . 75 

PANEL 26 

EFFECTS OF LAUNCH EXPOSURE ON REFRACTORY CONCRETE PANELS. 

SAUEREISEN CEMENTS COMPANY. 

11 

THESE GRAPHS SHOW AVERAGE THICKNESS LOSSES AND HAXJHUH PJT DEPTHS 

FOR INDIVIDUAL LAUNCH EXPOSURES. TYPICAL FEATURES ARE INDICATED 

AND JXPLAlNED FOR PANEL 2. FIGURE 1. Q.V. 

226 
\ 

MTB-2SO-R4 

OISTR~BUTION LIST 

OFFICE SYMBOL/NAME 

00-ME0-1/M. OLSEN 

SF-ENG/ROBERT MILLER 

SF-SEC-3/RONALD GILLETT 

SI-PR0-14 

ZICR6/VAN OUSEN 

- PRC-2401/J. WALKER 

PRC-121~/GARY KURTZ 

DE-HAO/C.W.HOPPESCH 

DE-GSO/~. ROWE 

DD-HED-33/J. F. MC INERNY 

Sl-ENG-1/WAYNE PARRIS 

NASA 1 GODDARD SPACE FLIGHT CENTER 

GREENBELT. MARYLAND 20771 

ATTN: MR. RICHARD HARRIOTT · CODE 313 

DE·HAO·Z 

NASA/TM-2010-216293R 

NO. OF COPIES 

5 

1 

1 

1 

1 

1 

1 

1 

1 

All THAT ARE 

LEFT 

PLEASE NOTIFY THIS OFFICE (867-4614/DE-HA0-2) IF THERE ARE ANY 

CHANGES IN NAHE 1 ADDRESS. OFFICE SYHBOL 1 ETC. • THANk YOU! 

227 
NASNTM-2010-216293R 

. This page intentionally left blank. 

228 
NASA/TM-2010-216293R 

APPENDIX H. 93-4436, STS-55 

229 
NASA/TM-2010-216293R 

NASA 

DIRECTOR OF ENGINEERING DEVELOPMENT 

DIRECTOR, MECHANICAL ENGINEERING 

MATERIALS SCIENCE LABORATORY 

FAILURE ANALYSIS AND MATERIALS EVALUATION BRANCH 

PHYSICAL TESTING SECTION 

DM-MSL-24, ROOM 1218, O&C BUILDING 

KENNEDY SPACE CENTER, FLORIDA 32899 

JULY 19, 1993 

REPORT NUMBE~~J-4436 

SUBJECT: EXPOSURE TEST OF REFRACTORY CONCRETE TEST SPECIMENS TO 

SOLID ROCKET BOOSTER (SRB) EXHAUST DURING LAUNCH OF 

STS-55, LC-39A 

PARTICIPANTS: J. GAY, NASA/DM-MSL-23 JOHN F. KENNEDY SPACE CENTER UBRARY 

H. KIM, NASA/DM-MSL-1 DOCUMENTS DEPARTMENT 

P. PETERSEN, NASA/ DM-MSL-23 REFERENCE COPY 

p. WELcH, NASA/DM-MsL-24 Ace 7''133 

RELATED DOCUMENTATION: 

1.0 FOREWORD 

KSC-SPEC-P-0012, REFRACTORY CONCRETE, SPECIFICATION 

FOR 

MTB-503-83, EXPOSURE TEST OF REFRACTORY CONCRETE 

TEST PANELS TO SRB EXHAUST DURING THE FIRST SEVEN 

STS LAUNCHES: SUMMARY REPORT 

MCB-34-93 AND MCB 429-93, MICROCHEMICAL ANALYSIS 

REPORTS 

1.1 The Florida Product Innovation Center, a state 

university system service program, requested that the 

NASA Failure Analysis and Materials Evaluation Branch 

evaluate a refractory concrete product developed by 

Mitec, Inc . , (Mitec) for potential applications such as 

the flame trenches at Kennedy Space Center (KSC). 

1.2 Mitec was provided with a copy of KSC-SPEC-P-0012, 

which contains the requirements for qualification test 

samples. 

1.3 Six refractory concrete test specimens, 75 ern square by 

10 em thick, were provided by Mitec for testing. 

Reference specimens of Fondu Fyre, the refractory 

material currently used at KSC, were fabricated by the 

Shuttle Processing Contractor (SPC). 

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2.0 TEST PROCEDURE 

2.1 Six test specimens, four Mitec and two Fondu Fyre, were selected for test. The thickness of each specimen was measured at 13 template defined locations. The test 

specimens were installed on the two test stands located 

at the north side base of the LC-39A flame deflector 

(see Figure 1). Each stand holds three test specimens. 

These locations allow the test specimens to be exposed 

to SRB exhaust blast during STS launches from the pad. 

2.1.1 Looking south at the flame deflector, the test 

locations are identified alphabetically from left 

to right as "A" through "F". 

FLAME DEFLECTOR NORTH SIDE 

VIEW LOOKING DOWN INTO FLAME PIT 

TO IDENTIFY TEST LOCATIONS "A" THROUGH "F" 

2.1.2 The reference, Fondu Fyre, test specimens were 

installed at Location "A" on the left side of the 

east stand, and at Location "E" in the center of 

the west stand. 

2.1.3 The Mitec test specimens were installed at the 

remaining four locations: "B," "C," "D," and "F." 

2.1.4 Samples from the test specimens were obtained for 

petrographic analysis. 

2.2 The installed refractory test specimens were subjected 

to the SRB exhaust blast during the launch of STS-55 on April 26, 1993 . 

2.3 After the launch, the test specimens were examined and 

documented in place, removed, and measured for thickness 

changes. 

2.3 . 1 The test specimens at Locations "A" and "B" were 

lost during the launch. The refractory material on the face of the flame deflector broke loose, 

destroying the two test specimens. · 

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2.3.2 The Mitec specimen at Location "C" (see Figures 2 

and 3) was grossly eroded and broke into numerous small pieces when removed from the stand. Post 

test thickness measurements were not attempted on this specimen. 

2.3.3 Half of the Mitec specimen at Location "D" (see 

Figures 4 and 5) was missing and broke into 

several pieces when removed from the stand. Post 

test thickness measurements were not attempted on this specimen. 

2.3.4 The Fondu Fyre reference test specimen at 

Location "E" appeared intact and was coated with a metallic material, except for the pitted areas. The test specimen was removed intact and post 

test thickness measurements were made of the 

major pieces. Chips of the metallic material were removed from the specimen to determine the 

layer thickness and identify the material. 

2.3.5 The Mitec specimen at Location "F" (see Figures 8 

and 9) appeared intact with only one noticeable 

crack traversing the face, and except for the 

pitted areas was coated with a metallic material. 

This specimen broke into seven major pieces when 

removed from the stand. Post-test thickness 

measurements were made of the major pieces. 

3.0 RESULTS 

Chips of the metallic material were removed from 

the specimen to determine the layer thickness and 

identify the material. 

The metrology data for the test specimens are presented 

below: 

Test Location "E" "F" 

Material Fondu Fyre Mitec 

Original 

Thickness (Avg.) 10.124 em 10.056 em 

Post Test 

Thickness (Avg.) 10.056 em 10.119 em 

Metallic Layer 

Thickness (Est.) 0.173 em 0.178 em 

Post Test Thickness 

(Est.) 9.883 em 9.941 em 

Thickness Loss (Est.) 0.241 em 0.111 em 

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4. 0 ANALYTICAL TESTS AND RESULTS 

4.1 The samples were analyzed by optical microscope, 

electron microprobe with energy dispersive spectrometry, 

and x-ray diffraction. 

4.2 The four Mitec samples were similar in composition and 

in petrographic features, and were described 

collectively. The samples were moderately hard and 

compact. They were composed of coarse aggregate, fine 

sand, and cement paste. 

4.2.1 The coarse aggregate appeared to be composed 

totally of white pumaceous material containing 

Si02 (alpha-cristoballite) and glass phase. 

4.2.2 The fine sand was composed mainly of light and 

fine grained Si-Al rich material. 

4.2.3 The cement paste was light gray in color and 

contained small amounts of unhydrated portland 

cement and was moderately carbonated throughout. 

The cement paste was composed of glass, 

unhydrated portland cement, ca5si60l6(0H)2 

(tobermorite), Al(OH) 3 , CaC03 (calc~te), 

Ca3Al2o6 , and Ca2 (Si03 )(0H) 2 • 

4.2.4 The Mitec samples were air entrained. The air 

content was estimated to be less than seven percent. 

4.3 The Fondu Fyre samples were also moderately hard and 

compact. They were composed of coarse aggregate, fine 

sand, and cement paste. 

4.3.1 The coarse aggregate appeared to be composed 

totally of dark coarse grained slag containing 

Fe2Ti04 (ulvospinel), CaFe3AlSi06 (essenite) , Ca(Fe,Mg)Si2o 6 (augite), glass, and 

ca5Al2 (si2Al2 )o10 (0H) 2 (margarite). 

4.3.2 The fine sand was composed mainly of light 

colored and fine grained Si-Al rich material. 

4.3.3 The cement paste was dark gray in color, 

contained prominent amounts of unhydrated 

portland cement, and was moderately carbonated 

throughout. The cement paste was composed of 

ca si~o 16 (0H) 2 (tobermorite), Al(OH) 3 , caco3 (calc~te), unhydrated portland cement, Ca3Al2o 6 , and amorphous materials. 

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93-4436 5 

4.3 . 4 The Fondu Fyre samples were also air entrained. 

The air content was estimated to be less than seven percent. 

4.4 The chips removed from the surface of the two refractory 

concrete test specimens were composed of Al203, 

(corundum). 

5.0 CONCLUSIONS 

5.1 Only two out of six test specimens survived the launch 

of STS-55. 

5.2 The poor performance of several Mitec refractory 

concrete test specimen resulted from their being 

fabricated without the steel mesh reinforcement as required in KSC-SPEC-P-0012. 

5.3 Mitec, Inc., has indicated that they will fabricate two more test specimens with steel mesh reinforcement. 

tNVBSTlGATOR' 

Pete1 

Y~;Jv!~ J. Welch 

APPROVAL:~~ N. j1Salvail, Chief, t?~~~~~~~--~~ Physical Testing Section -­

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FI GURE 1 

VIEW LOOKING SOUTH I N LC- 39A FLAME TRENCH AT REFRACTORY CONCRETE 

TEST STANDS AT BOTTOM OF FLAME DEFLECTOR. THE TEST STANDS ARE 

LOCATED BENEATH THE TWO SRB FLAME PITS IN THE MOBILE LAUNCH 

PLATFORMS. 

235 

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FIGURE 2 

MITEC TEST SPECIMEN AT LOCATION "C" PRIOR TO LAUNCH. NOTE THE 

TEST SPECIMENS ARE GROUTED WITH SILICONE RUBBER TYPE ABLATIVE 

MATERIAL. 

236 

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FIGURE 3 

MI TEC TEST SPECIMEN AT LOCATION "C" AFTER LAUNCH. THE TEST 

SPECIMEN CONTAINED NUMEROUS CRACKS AND WAS GROSSLY ERODED, 

ESPECIALLY ALONG THE CRACK LINES. 

237 

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FIGURE 4 

MITEC TEST SPECIMEN AT LOCATION "0" PRI OR TO LAUNCH. 

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FIGURE 5 

MITEC TEST SPECIMEN AT LOCATION "D" AFTER LAUNCH. OVER HALF OF 

THE TEST SPECIMEN WAS MISSING AND THE REMAINDER CONTAINED NUMEROUS 

CRACKS. AREAS OF HEAVY EROSION PITTING WERE NOTED ON THE FACE 

ADJACENT TO THE FLAME DEFLECTOR. THE SURFACE OF THE SPECIMEN 

APPEARED TO BE COATED WITH A ROUGH METALLIC LAYER EXCEPT IN THE 

AREAS OF THE PITTING. 

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FIGURE 6 

11 

. ;,:, ' ; .... ,.,.,,r 

FONOU FYRE TEST SPECIMEN AT LOCATION "E" PRIOR TO LAUNCH. 

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, . 

FIGURE 7 

FONDU FYRE TEST SPECIMEN AT LOCATION "E" AFTER LAUNCH. AREAS OF 

HEAVY EROSION PITTING WERE NOTED ON THE FACE ADJACENT TO THE FLAME 

DEFLECTOR. THE SURFACE OF THE SPECIMEN APPEARED TO BE COATED WITH 

A ROUGH METALLIC LAYER EXCEPT IN THE AREAS OF THE PITTING. 

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FIGURE 8 

MI TEC TEST SPECIMEN AT LOCATION "F" PRIOR TO LAUNCH. 

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FIGURE 9 

.~ .... -- ' 

·. '· · .. 

• 

14 

' . 

MITEC TEST SPECIMEN AT LOCATION "F" AFTER LAUNCH. SEVERAL CRACKS 

WERE EVIDENT IN THE FACE OF THE TEST SPECIMEN AND AREAS OF HEAVY 

EROSION PITTING WERE NOTED ON THE FACE ADJACENT TO THE FLAME 

DEFLECTOR. THE SURFACE OF THE SPECIMEN APPEARED TO BE COATED 

WITH A ROUGH METALLIC LAYER EXCEPT IN THE AREAS OF THE PITTING. 

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244 
REPORT DOCUMENTATION PAGE I Form Approved 

OMB No. 0704-0188 

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1. REPORT DATE (00-MM-YYYY) 12. REPORT TYPE 13. DATES COVERED (From- To) 

31-12-20110 Technical Memorandum 2009-2010 

4. TITLE AND SUBTITLE Sa. CONTRACT NUMBER 

Testing Requirements for Refractory Materials 

Sb. GRANT NUMBER 

Sc. PROGRAM ELEMENT NUMBER 

6. AUTHOR(S) Sd. PROJECT NUMBER 

(1) Calle·, Luz Marina (5) Kolody, Mark R. 

Se. TASK NUMBER 

(2) Hintze, Paul E. (6) Sampson, Jeffrey W. 

(3) Parlier, Christopher R. (7) Montgomery, Eliza M. Sf. WORK UNIT NUMBER 

(4) Curran, Jerome P. 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION 

REPORT NUMBER 

1,2, 3, 5, 7NE-L2-C 

Kennedy Space Center, Florida 32899 

4

' 

5 ASRC Aerospace 

Kennedy Space Center 32899 

9. SPONSOFUNG/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITOR'S ACRONYM(S) 

Nationall Aeronautics and Space Administration 

Washin1:ton, D.C. 20546-0001 

11 . SPONSORING/MONITORING 

REPORT NUMBER 

NASA/TM-2010-216293R 

12. DISTRIBilJTION/AVAILABILITY STATEMENT 

Unclassified- Unlimited Distribution: Standard 

Subject Category: 

Availability: NASA CASI (301) 621-0390 

13. SUPPLEIMENTARY NOTES 

This revision changes paragraph '10 days" to "14 days" in paragraph 3.2.9. 

14. ABSTRACT 

Launch Pads 39A and 39B currently use refractory material (Fondu Fyre) in the flame trenches. This material was initially approved for the 

Saturn prog~ am. This material had a lifetime of I 0 years according to the manufacturer, and it has been used for over 40 years. As a 

consequenc1!, the Fondu Fyre at Launch Complex 39 requires repair subsequent to almost every launch. A review of the literature indicates 

that the gunned Fondu Fyre refractory product (WA-1 G) was never tested prior to use. 

With the recent severe damage to the flame trenches, a new refractory material is sought to replace Fondu Fyre. In order to replace Fondu 

Fyre, a methodology to test and evaluate refractory products was developed. This paper outlines this methodology and discusses current 

testing requirements, as well as the laboratory testing that might be required. Furthermore, this report points out the necessity for subscale 

testing, the locations where this testing can be performed, and the parameters that will be necessary to qualify a product. The goal is to 

identify a more durable refractory material that has physical, chemical, and thermal properties suitable to withstand the harsh environment of 

the launch p·ads at KSC. 

1S. SUBJECT TERMS 

refractory materials, Fondu Fyre, flame trench 

17. LIMITATION OF 

16. SECURITY CLASSIFICATION OF: ABSTRACT 

uu a. REPORT ,b. ABSTRACT I c. THIS PAGE 

u u u 

18. NUMBER 

OF 

PAGES 

252 

19a. NAME OF RESPONSIBLE PERSON 

Dr. Luz Marina Calle 

19b. TELEPHONE NUMBER (Include area code) 

I (321l 867-3278 

Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std. Z39