Flight Motor Set 360L001 (STS-26R) Final Report

TWR-17272 - V_L-_

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Flight Motor Set 360L001(STS-26R)Final Report

j Volume V (Nozzle Component)

July 1989

Prepared for

National Aeronautics and Space Administration

George C. Marshall Space Flight Center

Marshall Space Flight Center, Alabama 35812

Contract No. NAS8-30490

DR No. 3-5WBS No. 4B601-03-08

ECS No. 956

SPACE OPERATIONS

CORPORA T/ON

PO Box 707, Bngham C_ty, UT 84302-0707 (801)863-351 l

PuOhcatlons No. 89435

_H(,HA_A-C_-l_3771) FLI_ T MIJTUR get 3{_n,LC'O]

(STS-_6,e). VOLUME 5: NOTZLL CQM_I)N_iNT Final

_{eport (Thioko| Corp.) _71 i; C::ICL 2IH

r, _ I ? n

NqO- 13 5 _:]

TVR-17272

360L001 (STS-26R)

Final Report

Nozzle Component

July 1989

Prepared by:

/ R._/ Gefrge/

NozzN_/TVg Design

S. Graves, Manager

Non- Me ta Ii ic Component s

#._ '_, {,t'}f.<.-,r__

R. B._/Roth, MaNager

N_ Work Center

System_Safety I.' 0

Re14_se

Approved by: _ _ _,

J. E. Fonnesbeck, Supervisor

Nozzle/TVC Design

--_r. L. Diehl

SRM _z..leyrograms

_CORPORA_ON

SPACE OPERATTONS

TABLE OF CONTENTS

i.0 INTRODUCTION .................................................... 1

2.0 OBJECTIVES ...................................................... 2

3.0 SUMMARY/CONCLUSIONS ............................................. 2

4.0 RESULTS/DISCUSSION .............................................. 3

4.1 STS-26A (L_) Nozzle/Flex Bearing ........................... 4

4.1.1 STS-26A Nozzle Components ........................... 4

4,1.2 STS-26A Nozzle Internal Joints ...................... 15

4.2 STS-26B Nozzle Flex Bearing ................................ 21

4.2.1 STS-26B Nozzle Components ........................... 214,2.2 STS-26B Nozzle Internal Joints ...................... 33

4.3 Instrumentation ............................................ 38

5.0 DISCREPANCY REPORTS AND PROCESS DEPARTURES ...................... 38

5.1 STS-26A DRs and PDs ........................................ 38

5.2 STS-26B DRs and PDs ........................................ 41

6.0 NOZZLE COMPONENT PROGRAM TEAM (NCPT) RECOMMENDATIONS AND

REDESIGN PROGRAM REVIEW BOARD (RPRB) ASSESSMENT ................. 44

6,1 STS-26A Nozzle ............................................. 44

6,2 STS-26B Nozz]e ............................................. 45

APPENDIX A ........................................................... A-I

APPENDIX B ........................................................... B-I

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LIST OF FIGURES

Figure I STS-26 Nozzle Components ..................................... 46

Figure 2a STS-26A Nozzle Material ...................................... 47

Figure 2b STS-26B Nozzle Material ...................................... 48

Figure 3 STS-26A Joint Flow Surface Gap Openings ...................... 49

Figure 4 STS-26A Forward Nozzle Assembly .............................. 50

Figure 5 STS-26A Forward Nozzle Assembly .............................. 51

Figure 6 STS-26A Forward Nozzle Assy (External) 0 o to 90 = to 180 °...52

Figure 7 STS-26A Forward Nozzle Assy (External) 180 o to 270 • to 0 °..53

Figure 8 STS-26A Forward Nozzle Assy (Internal) 0 o _ 90 o ............ 54

Figure 9 STS-26A Forward Nozzle Assy (Internal) 180 o to 270 o ........ 55

Figure 10 STS-26A Aft Exit Cone Fragment ............................... 56

Figure II STS-26A Aft Exit Cone Fragment ............................... 57

Figure 12 STS-26A Forward Exit Cone Dimpled Erosion (270 Degrees) ...... 59

Figure 13 STS-26A Forward Exit Cone BondlJne Separations ............... 60

Figure 14 STS-26A Forward Exit Cone Liner Section (0 Degrees) .......... 61

Figure 15 STS-26A Forward Exit Cone Liner Section (90 Degrees) ......... 62

Figure 16 STS-26A Forward Exit Cone Liner Section (180 Degrees) ........ 63

Figure 17 STS-26A Forward Exit Cone Liner Section (270 Degrees) ........ 64

Figure 18 Forward Exit Cone Assembly Nozzle Stations ................... 66

Figure 19 STS-26A Throat Ring Impact Marks (130 Degrees) ............... 67

Figure 20 STS-26A Throat Assembly Bondline Separations ................. 68

Figure 21 STS-26A Throat/Throat Inlet Section (0 Degrees) .............. 69

Figure 22 STS-26A Throat/Throat Inlet Section (90 Degrees) ............. 70

Figure 23 STS-26A Throat/Throat Inlet Section (180 Degrees) ............ 71

Figure 24 STS-26A Throat/Throat Inlet Section (270 Degrees) ............ 72

Figure 25 Throat Inlet Assembly Erosion Measurement Stations ........... 74

Figure 26 STS-26A -503 Ring Impact Marks (3 Degrees) ................... 75

Figure 27 STS-26A -504 Rlng Impact Mark (85 Degrees) ................... 76

Figt_re 28 STS-26A Nose Cap Wedgeout (20 Degrees) ....................... 77

Figure 29 STS-26A Nose Inlet Assembly Bondline Separations ............. 78

Figure 30 STS-26A Forward Nose Ring and Aft Inlet Ring (-503 and -504)

Section (0 Degrees) .......................................... 79

Figure 31STS-26A Forward Nose Ring and Aft Inlet Ring (-503 and -504)

Section (90 Degrees) ......................................... 80


Section (180 Degrees) ........................................ 81


Section (270 Degrees) ........................................ 82

Figure 34 STS-26A Nose Cap Section (0 Degrees) ......................... 83

Figure 35 STS-26A Nose Cap Section (90 Degrees) ........................ 84

Figure 36 STS-26A Nose Cap Section (180 Degrees) ....................... 85

Figure 37 STS-26A Nose Cap Section (270 Degrees) ....................... 86

Figure 38 Nose Inlet Assembly Erosion Measurement Stations ............. 90

Figure 39 STS-26A Nose Inlet Housing Bonding Surfaces (0 Degrees) ...... 91

Figure 40 STS-26A Nose Inlet Housing Bonding Surfaces (90 Degrees) ..... 92

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LIST OF FIGURES (continued)



Figure 43 STS-26A Nose Inlet Housing (Nose Cap) Bonding Surface

(90 Degrees) .................................................. 95

Figure 44 STS-26A Nose Inlet Housing (Nose Cap) Bonding Surface

(180 Degrees) ................................................. 96

Figure 45 STS-26A Nose Inlet Housing (-504 Ring) Bonding Surface

(90 Degrees) .................................................. 97

Figure 46 STS-26A Nose Inlet Housing (-504 Ring) Bonding Surface

(270 Degrees) ................................................. 98

Figure 47 STS-26A CowI/OBR Closeup (0 Degrees) .......................... 99

Figure 48 STS-26A Cowl/OBR Closeup (160 Degrees) ........................ 100

Figure 49 STS-26A Cowl/OBR Closeup (180 Degrees) ....................... I01

Figure 50 STS-26A Cowl/OBR Closeup (320 Degrees) ....................... 102

Figure 51STS-26A Cowl Ring Section (0 Degrees) ........................ 103

Figure 52 STS-26A Cowl (90 Degrees) .................................... 104

Figure 53 STS-26A Cowl (180 Degrees) ................................... 105

Figure 54 STS-26A Cowl (270 Degrees) ................................... 106

Figure 55 Cowl Ring and Outer Boot Ring Erosion Measurement Stations ...III

Figure 56 STS-26A OBR Aft End Delaminations (260 Degrees) .............. 112

Figure 57 STS-26A Fractured OBR Aft Tip Adjacent to Flex Boot .......... 113

Figure 58 STS-26A Outer Boot Ring Section (0 Degrees) .................. 114

Figure 59 STS-26A Outer Boot Ring/Flex Boot (90 Degrees) ............... 115

Figure 60 STS-26A Outer Boot Ring/Flex Boot (180 Degrees) .............. 116

Figure 61STS-26A Outer Boot Ring/Flex Boot (280 Degrees) .............. 117

Figure 62 STS-26A Flex Boot (Cavity Side - 0 Degrees) .................. 118

Figure 63 STS-26A Flex Boot (Cavity Side - 120 Degrees) ................ 119

Figure 64 STS-26A Flex Boot (Cavity Side - 240 Degrees) ................ 120

Figure 65 STS-26A Fixed Housing Wedgeout (280 Degrees) ................. 122

Figure 66 STS-26A Fixed Housing Section (0 Degrees) .................... 123

Figure 67 STS-26A Fixed Housing Section (90 Degrees) ................... 124

Figure 68 STS-26A Fixed Housing Section (180 Degrees) .................. 125

Figure 69 STS-26A Fixed Housing Section (270 Degrees) .................. 126

Figure 70 Fixed Housing Liner Erosion Measurement Station .............. 128

Figure 71STS-26A Bearing Protector (0 Degrees) ........................ 129

Figure 72 STS-26A Bearing Protector (120 Degrees) ...................... 130

Figure 73 STS-26A Bearing Protector (240 Degrees) ...................... 131

Figure 74 STS-26A Flex Bearing (0 Degrees) ............................. 132

Figure 75 STS-26A Flex Bearing (180 Degrees) ........................... 133

Figure 76 STS-26A Forward Exit Cone-to-Aft Exit Cone Joint Interface ...134

Figure 77 STS-26A Aft Exit Cone Forward End (0 Degrees) ................ 135

Figure 78 STS-26A Aft Exit Cone Forward End (120 Degrees) .............. 136

Figure 79 STS-26A Aft Exit Cone Forward End (240 Degrees) .............. 137

Figure 80 STS-26A Forward Exit Cone - Aft End (0 Degrees) .............. 138

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Figure 81STS-26A Forward Exit Cone - Aft End (120 Degrees) ............ 139

Figure 82 STS-26A Forward Exit Cone - Aft End (240 Degrees) ............ 140

Figure 83 STS-26A Forward Exit Cone Aft End Black Corrosion

(338 Degrees) ................................................ 141

Figure 84 STS-26A Forward Exit Cone Aft End White Corrosion

(45 Degrees) ................................................. 142

Figure 85 STS-26A Forward Exit Cone Aft End Scratch Mark (90 Degrees) ..143

Figure 86 STS-26A Throat/Forward Exit Cone Joint ....................... 144

Figure 87 STS-26A Forward Exit Cone - Forward End (0 Degrees) .......... 145

Figure 88 STS-26A Forward Exit Cone - Forward End (120 Degrees) ........ 146

Figure 89 STS-26A Forward Exit Cone - Forward End (240 Degrees) ........ 147

Figure 90 STS-26A Throat Aft End (0 Degrees) ........................... 148

Figure 91STS-26A Throat Aft End (120 Degrees) ......................... 149

Figure 92 STS-26A Throat Aft End (240 Degrees) ......................... 150

Figure 93 STS-26A Throat Aft End - Blowpath (310 Degrees) .............. 151

Figure 94 STS-26A Nose Inlet/Throat Housing Joint ...................... 152

Figure 95 STS-26A Throat Forward End (0 Degrees) ....................... 153

Figure 96 STS-26A Throat Forward End (120 Degrees) ..................... 154

Figure 97 STS-26A Throat Forward End (240 Degrees) ..................... 155

Figure 98 STS-26A Aft Inlet (-504) Ring Aft End (0 Degrees) ............ 156

Figure 99 STS-26A Aft Inlet (-504) Ring Aft End (120 Degrees) .......... 157

Figure I00 STS-26A Aft Inlet (-504) Ring Aft End (240 Degrees) ......... 158

Figure I01STS-26A Throat Forward End Blowpath (136 Degrees) ........... 159

Figure 102 STS-26A Aft Inlet (-504) Ring Aft End Blowpath (136 Degrees).160

Figure 103 STS-26A Nose Inlet Housing/Flex Bearing Joint ............... 161

Figure 104 STS-26A Cowl Forward End (0 Degrees) ........................ 162

Figure 105 STS-26A Cowl Forward End (120 Degrees) ...................... 163

Figure 106 STS-26A Cowl Forward End (240 Degrees) ...................... 164

Figure 107 STS-26A Nose Cap Aft End (0 Degrees) ........................ 165

Figure I08 STS-26A Nose Cap - Aft End (120 Degrees) .................... 166

Figure 109 STS-26A Nose Cap - Aft End (240 Degrees) .................... 167

Figure II0 STS-26A Bearing Forward End Ring (0 Degrees) ................ 168

Figure iii STS-26A Bearing Forward End Ring (120 Degrees) .............. 169

Figure 112 STS-26A Bearing Forward End Ring (240 Degrees) .............. 170

Figure ll3 STS-26A Cowl Forward End - Blowpath Location (216 Degrees) ..171

Figure 114 STS-26A Flex Bearing/Fixed Housing Joint .................... 172

Figure 115 STS-26A Fixed Housing Forward End (0 Degrees) ............... 173

Figure 116 STS-26A Fixed Housing Forward End (120 Degrees) ............. 174

Fig,lre 117 STS-26A Fixed Housing Forward End (240 Degrees) ............. 175

Figure I18 STS-26A Bearing A_t End Ring (0 Degrees) .................... 176

Figure 119 STS-26A Bearing Aft End Ring (120 Degrees) .................. 177

Figure 120 STS-26A Bearing Aft End Ring (240 Degrees) .................. 178

Figure 121STS-26A Bearing Aft End Ring (300 Degrees) .................. 179

Figure 122 STS-26A Fixed Housing Forward End Rust on Metal Surfaces

(15 Degrees) ................................................ 180

Figure 123 STS-26A Bearing Aft End Ring - White Corrosion Spot

(260 Degrees) ............................................... 181

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Figure 124 STS-26B Joint Flow Surface Gap Openings ..................... 182

Figure 125 STS-26B Forward Nozzle Assembly ............................. 183

Figure 126 STS-26B Forward Nozzle Assembly ............................. 184

Figure 127 STS-26B Forward Nozzle Assembly (External) 0 o to 90 oto 180 o .................................................... 185

Figure 128 STS-26B Forward Nozzle Assembly (External) 180 o to 270 oto 0 o ...................................................... 186

Figure 129 STS-26B Forward Nozzle Assembly (Internal) 0 o to 90 oto 180 o .................................................... 187

Figure 130 STS-26B Forward Nozzle Assembly (Internal) 180 o to 270 oto 0 o ...................................................... 188

Figure 131STS-26B Aft Exit Cone Fragment .............................. 189

Figure 132 STS-26B Forward Exit Cone Dimpled Erosion ................... 191

Figure 133 STS-26B Forward Exit Cone Bondline Separations .............. 192

Figure 134 STS-26B Forward Exit Cone Liner Section (0 Degrees) ......... 193

Figure 135 STS-26B Forward Exit Cone Liner Section (90 Degrees) ........ 194

Figure 136 STS-26B Forward Exit Cone Liner Section (180 Degrees) ....... 195

Figure 137 STS-26B Forward Exit Cone Liner Section (270 Degrees) ....... 196

Figure 138 STS-26B Forward Exit Cone Corrosion (0 Degrees) ............. 198

Figure 139 STS-26B Forward Exit Cone Corrosion (80 Degrees) ............ 199

Figure 140 STS-26B Forward Exit Cone Corrosion (30 Degrees) ............ 200

Figure 141STS-26B Forward Exit Cone Corrosion (60 Degrees) ............ 201

Figure 142 STS-26B Forward Exit Cone Corrosion (120 Degrees) ........... 202



Figure 145 STS-26B Throat Inlet Ring Wedgeout (28 - 40 Degrees) ........ 205

Figure 146 STS-26B Throat Assembly Bondline Separations ................ 206

Figure 147 STS-26B Throat/Throat Inlet Sectiun (0 Degrees) ............. 207

Figure 148 STS-26B Throat/Throat Inlet Section (180 Degrees) ........... 208

Figure 149 STS-26B Throat/Throat Inlet Section (270 Degrees) ........... 209

Figure 150 STS-26B (-503) Ring Impact Marks (125 Degrees) .............. 211

Figure 151STS-26B (-503) Ring Impact Marks (185 Degrees) .............. 212

Figure 152 STS-26B Typical Nose Cap Aft End Wedgeout (Post-Burn)

(170 Degrees) ....................................... 213

Figure 153 STS-26B Nose Inlet Assembly Bondline Separations ............ 214

Figure 154 STS-26B Forward Nose Ring and Aft Inlet Ring (-503 and -504)

Section (0 Degrees) ................................. 215


Section (90 Degrees) ................................ 216


Section (180 Degrees) ............................... 217


Section (270 Degrees) ............................... 218

Figure 158 STS-26B Nose Cap Section (0 Degrees) ........................ 219

Figure 159 STS-26B Nose Cap Section (90 Degrees) ....................... 220

Figure 160 STS-26B Nose Cap Section (180 Degrees) ...................... 221

Fig,are 161STS-26B Nose Cap Section (270 Degrees) ...................... 222

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Figure 162 STS-26B Nose Inlet Housing Bonding Surfaces (0 Degrees) ..... 226

Figure 163 STS-26B Nose Inlet Housing Bonding Surfaces (90 Degrees) .... 227Figure 164 STS-26B Nose Inlet Housing Bonding Surfaces (180 Degrees) ...228

Figure 165 STS-26B Nose Inlet Housing Bonding Surfaces (270 Degrees) ...229Figure 166 STS-26B Nose Cap Bonding Surface (0 Degrees) ................ 230

Figure 167 STS-26B Nose Cap Bonding Surface (180 Degrees) .............. 231

Figure 168 STS-26B Nose Inlet Housing Pitting (180 Degrees) ............ 232Figure 169 STS-26B Cowl/OBR Closeup (0 Degrees) ........................ 233

Figure 170 STS-26B Cowl/OBR Closeup (180 Degrees) ...................... 234Figure 171STS-26B Cowl/OBR Closeup (320 Degrees) ...................... 235

Fig.re 172 STS-26B Cowl Vent Hole Plugged with Slag (160 Degrees) ...... 236I;'ig. le 11_ STS 26B (:owl Ring Wedg(-out Containing Slag

(120 - 137 Degrees) ......................................... 23]Figure 174 STS-26B Cowl Ring Section (0 Degrees) ....................... 238

Figure 175 STS-26B Cowl Ring Section (45 Degrees) ...................... 239Figure 176 STS-26B Cowl Ring Section (180 Degrees) ..................... 240

Figure 177 STS-26B Cowl Ring Section (270 Degrees) ..................... 241Figure 178 STS-26B Outer Boot Ring Aft Tip Delamination (0 Degrees) .... 247

Figure 179 STS-26B Outer Boot Ring Section (0 Degrees) ................. 248

Figure 180 STS-26B Outer Boot Ring Section (100 Degrees) ............... 249Figure 181STS-26B Outer Boot Ring Section (160 Degrees) ............... 250

Figure 182 STS-26B Outer Boot Ring Section (280 Degrees) ............... 251

Figure 183 STS-26B Flex Boot (Cavity Side - 0 Degrees) ................. 252Figure 184 STS-26B Flex Boot (Cavity Side - 120 Degrees) ............... 253

Figure 185 STS-26B Flex Boot (Cavity Side - 240 Degrees) ............... 254Figure 186 STS-26B Fixed Housing Section (0 Degrees) ................... 256Figure 187 STS-26B Fixed Housing Section (90 Degrees) .................. 257

Figure 188 STS-26B Fixed Housing Section (180 Degrees) ................. 258Figure 189 STS-26B Fixed Housing Section (270 Degrees) ................. 259Fig.re 190 STS-26B Bearing Protector (O Degrees) ....................... 261}"ig_le 191STS-26B Bearing Protector (]20 Degrees) ..................... 262

|,'ig.[e 192 STS 26B Beai ing Pxotectot: (240 Degrees) ..................... 261_

Figure 193 STS-26B Flex Bearing (330 Degrees) .......................... 265Figure 194 STS-26B Aft Exit Cone-to-Forward Exit Cone Joint Interface ..266Figure 195 STS-26B Aft Exit Cone Forward End (0 Degrees) ............... 267

Fig,ire 196 STS-26B Aft Exit Cone Forward End (120 Degrees) ............. 268Figure 197 STS-26B Aft Exit Cone Forward End (240 Degrees) ............. 269

Figure 19B STS-26B Forward Exit Cone - Aft End (0 Degrees) ............. 270

Figure 199 STS-26B Forward Exit Cone - Aft End (120 Degrees) ........... 271

Figure 200 STS-26B Forward Exit Cone - Aft End (240 Degrees) ........... 272

Figure 201STS-26B Forward Exit Cone Aft End Corrosion ................. 273

Figure 202 STS-26B Forward Exit Cone Aft End Scuff Mark ................ 274

Figure 203 STS-26B Throat/Forward Exit Cone Joint ...................... 275

Figure 204 STS-26B Forward Exit Cone - Forward End (0 Degrees) ......... 276

Figure 205 STS-26B Forward Exit Cone - Forward End (120 Degrees) ....... 277

Figure 206 STS-26B Forward Exit Cone - Forward End (240 Degrees) ....... 278

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Figure 207 STS-26B Throat Aft End (0 Degrees) .......................... 279

Figure 208 STS-26B Throat Aft End (120 Degrees) ........................ 280

Figure 209 STS-26B Throat Aft End (240 Degrees) ........................ 281

Figure 210 STS-26B Nose Inlet/Throat Housing Joint ..................... 282

Figure 211STS-26B Throat - Forward End (0 Degrees) .................... 283

Figure 212 STS-26B Throat - Forward End (120 Degrees) ................... 284

Figure 213 STS-26B Throat - Forward End (240 Degrees) .................. 285

Figure 214 STS-26B Aft Inlet (-504) Ring - Aft End (0 Degrees) ......... 286

Figure 215 STS-26B Aft Inlet (-504) Ring - Aft End (120 Degrees) ....... 287

Figure 216 STS-26B Aft Inlet (-504) Ring - Aft End (240 Degrees) ....... 288

Figure 217 STS-26B Nose Inlet Houslng/Flex Bearing Joint ............... 289

Figure 218 STS-26B Cowl - Forward End (0 Degrees) ...................... 290

Figure 219 STS-26B Cowl - Forward End (120 Degrees) .................... 291

Figure 220 STS-26B Cowl - Forward End (240 Degrees) .................... 292

Figure 221STS-26B Nose Cap - Aft End (O Degrees) ...................... 293

Figure 222 STS-26B Nose Cap - Aft End (120 Degrees) .................... 294


Figure 224 STS-26B Bearing Forward End Ring (0 Degrees) ................ 296

Figure 225 STS-26B Bearing Forward End Ring (150 Degrees) .............. 297

Figure 226 STS-26B Bearing Forward End Ring (240 Degrees) .............. 298


Figure 228 STS-26B Cowl Forward End - Blowpath Location (18 Degrees) ...300

Figure 229 STS-26B Flex Bearing/Fixed Housing Joint .................... 301

Figure 230 STS-26B Bearing Aft End Ring (0 Degrees) .................... 302

Figure 231STS-26B Bearing Aft End Ring (120 Degrees) .................. 303

Figure 232 STS-26B Bearing Aft End Ring (240 Degrees) .................. 304

Figure 233 STS-26B Fixed Housing Forward End (0 Degrees) ............... 305

Figure 234 STS-26B Fixed Housing Forward End (120 Degrees) ............. 306

Figure 235 STS-26B Fixed Housing Forward End (240 Degrees) ............. 307

Figure 236 STS-26B Aft Exit Cone LDI Location (45 Degrees) ............. 308

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LIST OFTABLES

Table 1

Table 2Table 3

Table 4Table 5

Table 6

Table 7Table 8

Table 9

STS-26A Aft Exit Cone Post-Flight Polysulfide Groove RadialWidths ........................................................ 58STS 26A Forward Exit Cone Erosion and Char Data ............... 65

STS-26A Throat Assembly Erosion and Char Data ................. 73STS-26A Nose Inlet Rings (-503, -504) Erosion and Char Data ... 87

STS-26A Nose Cap Assembly Erosion and Char Data ............... 88STS-26A Cowl/OBR Erosion and Char Data ........................ 107

STS-26A Flex Boot Data Performance Margins of Safety .......... 121

STS-26A Fixed Housing Insulation Erosion and Char Data ........ 127STS-26B Aft Exit Cone Post-Flight Polysulfide Groove RadialWidths ........................................................ 190

Table 10 STS-26B Forward Exit Cone Erosion and Char Data ............... 197

Table 11STS-26B Throat Assembly Erosion and Char Data ................. 210Table 12 STS-26B Nose Inlet Rings (-503, -504) Erosion and Char Data ...223Table 13 STS-26B Nose Cap Assembly Erosion and Char Data ............... 224Table 14 STS-26B Cowl/OBR Erosion and Char Data ........................ 225

Table 15 STS-26B Flex Boot Data Performance Margins of Safety .......... 255

Table 16 STS-26B Fixed Housing Insulation Erosion and Char Data ........ 260

Table 17 STS-26B Maximum Bearing Protector Erosion in Line with CowlVent Holes .................................................... 264

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1.0 INTRODUCTION

A review of the performance and post-flight condition of the STS-26

Redesigned Solid Rocket Motor (RSRM) nozzles is presented in this document.

Applicable Discrepancy Reports (DRs) and Process Departures (PDs) are

presented in Section 5.0. The Nozzle Component Program Team (NCPT)

performance evaluation and the Redesign Program Review Board (RPRB)

assessment is included in Section 6.0.

The STS-26 nozzle assemblies were flown on the RSRM First Flight (Space

Shuttle Discovery) on 29 September 1988. The nozzles were a partially

submerged convergent/divergent movable design with an aft pivot point

flexible bearing. The nozzle assembly (Figure 1) incorporated the

following features:

a. RSRM forward exit cone with snubbers

b. RSRM fixed housing

c. Structural backup Outer Boot Ring (OBR)

d. RSRM cowl ring

e. RSRM nose inlet assembly

f. RSRM throat assembly

g. RSRM forward [lose and aft inlet ring

h. RSRM aft exit cone assembly with Linear-Shaped Charge (LSC)

i. RTV backfill in joints 1, 3, and 4

j. Use of EA913 NA adhesive in place of EA913 adhesive

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k. Redesigned nozzle plug

I. Carbon Cloth Phenolic (CCP) with 750 ppm sodium content

Figures 2a and 2b show the CCP material usage for the STS-26 forward nozzle

assemblies and aft exit cone assemblies.

2.0 OBJECTIVES

The RSRM First Flight test objectives, as outlined in TUR-17535 (MTI

Engineering Requirements Document for RSRM First Flight), are as follows

(CPW1-3600 paragraph numbers are in parentheses):

K. Demonstrate flex bearing system reusability (3.2.1.9.c).

Y. Post-flight inspection of flex bearing to determine sealing

performance in the flight environment (3.2.1.2.3.b).

Z. Post-flight inspection to verify no gas leaks occurred between

the flex bearing internal components (3.2.1.2.3.d).

AD. Post-flight inspection for flex bearing damage due to water

impact (3.2.1.4.6.a).

AE. Post-flight inspection to verify nozzle liner performance

(3.2.1.4.13).

AV. Post-flight inspection to verify remaining nozzle ablative

thicknesses (3.3.6.1.2.7).

Post-flight inspection to verify nozzle safety factors

(3.3.6.1.2.8).

3.0 SUMMARY/CONCLUSI ONS

Compliance to the objectives is discussed below.

K. Evaluation indicates no condition which would advernely at[ectthe reusability of the flex bearing system. Both flex bearings

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have met all of the refurbishment requirements and areacceptable for reuse.

Y. Preliminary inspection shows the flex bearings remained sealed

throughout all motor induced environments. Tensile leak tests

done during the refurbishment cycle indicated no leakage.

Z. Preliminary inspection shows the flex bearings maintained a

positive seal within the internal components. Tensile leak

tests done during the refurbishment cycle also indicated no

leakage.

AD. Both flex bearings have met all of the refurbishment

requirements indicating there was no damage due to water

impact.

AE. Evaluation of both nozzle liners revealed erosion profiles

similar to what has been observed on RSRM static test nozzles.

gedgeouts in the aft ends of the RH cowl (120 to 137 degrees)

and nose cap (5 to 20 degrees) contained small amounts of slag.

Sectioning of the liners showed that the wedgeouts occurred

post-burn.

AV. Measurements of the nozzle remaining ablative liner thicknesses

show that the design safety factors have been met.

Sectioning and measurement of the liners show that the

performance margins of safety are all positive.

4.0 RESULTS/DISCUSSION

All STS-26 post-flight nozzle observations are discussed in detail below.

CCP liner Performance Margins of Safety (PMS) are presented using measured

erosion, and corresponding measured char vah, es adjusted to the end of

action time.

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4.1STS-26A (LH) Nozzle/Flex Bearin G

Overall erosion of the STS-26A forward nozzle assembly CCP ablative liner

was smooth and uniform. All CCP delamlnations, wedgeouts, and pop-ups were

determined to be post-burn occurrences resulting from cooldown of the

liners. Blowpaths were observed in joints I, 2, and 4, but there was no

blowby, erosion, or heat effect to the primary O-rings. Small amounts of

corrosion were found on the metal surfaces of joints i, 3, 4, and 5, but no

pitting was observed. Heavy corrosion and pitting was found on the nose

inlet housing bonding surfaces when the phenolics were washed off.

Post-flight subassembly flow surface gaps are shown in Figure 3. Overall

views of the nozzle are shown in Figures 4 through 9.

4.1.1STS-26A Nozzle Components

STS-26A Aft Exit Cone Assembly

Overall views of the STS-26A aft exit cone fragment are shown in Figures I0

and II.

The aft exit cone was severed aft of the compliance ring by the LSC. The

nozzle severance system performance was nominal. The exit cone cut was

clean, with no unusual tearing or breaking. The remaining aft exit cone

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fragment showed missing CCP liner 360 degrees circumferentially. This is a

typical post-flight observation and occurs at LSC firing and at splashdown.

Glass Cloth Phenolic (GCP) plies exposed by the missing liner showed no

signs of heat effect.

The polysulfide groove fill on the forward end of the aft exit cone shoved

no separations. Post-flight measurements of the polysulfide groove radial

width (Table 1) show that the GCP insulator did not pull away from the

aluminum shell during cooldown. The polysulfide shrank axially aft up to

0.12 in.

There were no separations observed within the GCP insulator on the forward

end.

STS-26A Forward Exit Cone Assembly

Overall views of the STS-26A forward exit cone are shown in Figures 8 and

9.

The forward exit cone showed missing CCP liner over the center portion of

the cone 360 degrees circumferentially. This is a typical post-flight

observation and occurs at splashdown and during Diver Operated Plug (DOP)

insertion. The GaP insulator exposed by the missing liner showed no signs

of heat effect. The CCP liner remained bonded on the forward 11 inches and

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on the aft 8 inches of the cone. These portions showed nominal erosion

with no major washing or pocketing. The aft 8 inches of the liner showed

the typical dimpled erosion pattern that has occurred on all flight and

static test forward exit cones (see Figure 12). The maximum radial depth

of the dimpled erosion was 0.15 inch.

The aft end of the forward exit cone showed bondline separations between

the EA946 adhesive and the steel housing from 30 to 60 degrees and from 124

to 148 degrees. The maximum radial width of the separations was 0.025

inch. The forward end o£ the forward exit cone showed bondline separations

between the GCP and CCP (0.04 inch maximum radial width), and cohesive

separations within the GCP (0.04 inch maximum radial width) intermittently

around the circumference. Figure 13 lists the location and radial width

measurements of all separations on the forward exit cone. These

separations are typical observations which have been seen on previous

static test and flight nozzles, and have been shown to occur post-burn.

Photographs of the sectioned forward exit cone liner are presented in

Figures 14 through 17. Char and erosion analysis of the sections is

presented in Table 2. Figure 18 shows the location of the measurement

stations. All margins of safety were positive, with a minimum of 0.01

occurring at station 28 (180 degrees).

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STS-26A Throat Assembly

Overall views of the STS-26A throat assembly (throat ring and throat inlet

ring) are shown in Figures 8 and 9.

The post-fired mean diameter of the throat was 55.922 inches (erosion rate

of 8.42 mils/second based on an action time of 122.4 seconds). Nozzle

post-burn throat diameters have ranged from 55,787 to 56.38 inches. The

flow surface bondltne gap between the throat and throat inlet rings was

0.08 inch and is typical of past static test and flight nozzles.

Erosion of the throat and throat inlet rings was smooth and uniform with no

wedgeouts observed. Popped-up charred CCP material was observed on the

forward 1.5 inches of the throat ring at 10, 70, 210, 285, and 345 degrees.

Sharp edges indicate the popped-up material occurred after motor operation.

Impact marks were noted on the throat inlet ring and on the aft end of the

throat ring intermittently around the circumference. The largest was

located on the throat inlet ring at 130 degrees and measured 1 inch

circumferentially by 0.5 inch axially by 0.25 inch radially (Figure 19).

These marks most likely resulted from the loose aft and forward exit cone

CCP material inside the motor at splashdown.

Bondline separations between the EA913 NA adhesive and the steel throat

support housing were observed on the aft end circumferentially except from

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0 to 25 degrees and at 335 degrees. The maximum radial width of these

separations was O.10 inch. Metal-to-adhesive bondline separations

measuring 0.03 inch wide radially were observed on the forward end of the

throat assembly clrcumferentlally except at 180 degrees. Separations

between the adhesive and GCP and within the adhesive were observed at 110,

and 180 degrees, respectively. These also measured 0.03 inch wide

radlally. Figure 20 lists the location and radial width measurements of

all separations on the throat assembly. These separations are typical

observations which have been seen on previous static test and flight

nozzles, and have been shown to occur post-burn.

Photographs of the sectioned throat assembly liner are presented in Figures

21 through 24. Char and erosion analysis of the sections is presented in

Table 3. Figure 25 shows the location of the measurement stations. All

margins of safety were positive, with a minimum of 0.06 occurring at

station 8 (0 degrees).

STS-26A Nose Inlet Assembly

Overall views of the STS-26A nose inlet asssembly (forward nose ring, aft

inlet ring, and nose cap) are shown in Figures 4 and 5.

The ply angle of the forward nose (-503) ring was checked and found to be

of the RSRM design. The flow surface bondline gap between the -503 ring

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and the aft Inlet (-504) ring was 0.15 inch. The flow surface bondline gap

between the -503 ring and the nose cap was 0.05 inch. These post-fired

measurements are typical of past static test and flight nozzles.

The

or

on

These marks most likely resulted from the loose

CCP material inside the motor at splashdown.

-503 and -504 rings showed smooth erosion with no pockets, wash areas,

wedgeouts. Impact marks occurring after motor operation were observed

both rings intermittently around the circumference (Figures 26 and 27).

aft and forward exit cone

The nose cap showed smooth erosion with no pockets or major washes

observed. The aft 2 to 3 inches showed popped-up charred CCP material at

137, 280, 310, and 332 degrees. Typical post-burn wedgeouts on the aft 2

to 3 inches (Figure 28) were noted from 14 to 26, 40 to 93, 110 to 122, 156

to 172, and 248 to 265 degrees. The maximum radial depth was 0.5 inch at

the cowl interface. Sharp edges indicate the popped-up material and the

wedgeouts occurred after motor operation. No wedgeouts were observed on

the forward end of the nose cap.

The aft end of the nose inlet assembly (-504 ring aft end) showed metal to

adhesive bondline separations (0.01 inch maximum radial width) occurring

intermittently around the circumference. Bondline separations between the

EA946 adhesive and the GCF (0.01 inch maximum radial width) were also

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observed. Bondltne separations were observed on the aft end of the nose

cap between the metal and EA946 adhesive, and the adhesive and GCP

intermittently around the circumference. The maximum radial width of these

separations was 0.005 inch. One separation, 0.003 inch vide radially, was

noted within the adhesive from 26 to 28 degrees. Figure 29 lists the

location and radial width measurements of all separations on the nose inlet

assembly. These separations are typical observations seen on previous

static test and flight nozzles and have been shown to occur post-burn.

Photographs of the sectioned nose inlet assembly rings are presented in


presented in Tables 4 and 5. Figure 38 shows the location of the

measurement stations. All margins of safety were positive, with a minimum

of 0.05 occurring at station 39.5 (180 degrees) for the -503/-504 rings,

and 0.04 occurring at station 24 (90 degrees) for the nose cap.

Following the washout of the phenolics, it was found that the aluminum nose

inlet housing had extensive corrosion and pitting on all bonding surfaces

360 degrees circumferentially (Figures 39 through 46). This was also

found on the STS-26B (RB) nose inlet housing. The cause of this corrosion

has been attributed to seawater which enters bondline separations during

splashdown and retrieval (Ref. Memo L231-FY89-M130). The metal bonding

surfaces were not accessible until phenolic washout at Clearfield

Operations. Therefore, corrosion protection was not applied to these

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surfaces until approximately 4 months after flight. This hardware will be

inspected during refurbishment for compliance to ST_7-3434 (Refurbishment

Of And Acceptance Criteria For Space Shuttle SRM Nozzle Metal Hardware).

STS-26A Cowl Ring

Overall views of

Close-up views are shown in

appeared plugged with slag

Figure 48).

the STS-26A cowl ring are shown in Figures 6 and 7.

Figures 47 through 50. All cowl vent holes

on the Outer Diameter (OD) of the ring (see

Typical ridged erosion was observed intermittently around the cowl

circumference. The forward portion of the ring eroded a maximum of 0.15

inch greater than on the aft portion of the ring (Figure 47). This is a

result of the low ply angle of the cowl ring and has been observed on the

majority of flight and static test nozzles. There were no wedgeouts

observed on the cowl ring.

There were no bondline separations on the forward end of the cowl ring.

Photographs of the sectioned cowl ring are presented in Figures 51 through

54. Typical subsurface ply lifting was observed intermittently around the

circumference along the forward 2 inches of the cowl.

separation was 0.10 inch at 0 degrees (Figure 51).

of flow or erosion within the delaminations.

The largest ply lift

There was no evidence

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Char and erosion analysis of the sections is presented in Table 6 (Stations

0 through 7). Figure 55 shows the location of the measurement stations.

All margins of safety were positive, with a minimum of 0.19 occurring at


STS-26A Outer Boot Ring/Flex boot

Overall views of the STS-26A OBR and flex boot are shown in Figures 6 and

7. Close-up views of the OBR are shown in Figures 47 through 50. The

bondline between the OBR and cowl ring remained intact with no indications

of flow. The flow surface bondllne gap was 0.18 inch and is typical of

past static test and flight nozzles.

The structural backup 0BR was intact. The flow surfaces showed smooth

erosion with no pockets, major washes, or wedgeouts. Delaminations in the

charred CCP of the aft tip were observed 360 degrees circumferentially

(Figure 56). Charred CCP material on the aft tip adjacent to the flex boot

fractured and popped up over a majority of the circumference (Figure 57).

A large impact mark was located on the aft end of the OBR at 190 degrees

and measured 6 inches circumferentially (Figure 49). This may have been

due to the loose CCP material in the motor after splashdown. Sharp edges

on the surfaces indicate this occurred after motor operation.

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Photographs of the sectioned OBR are presented in Figures 58 through 61.

Char and erosion analysis of the sections is presented in Table 6 (Stations

8 through 12). Figure 55 shows the

All margins of safety were positive,


location of the measurement stations.

with a minimum of 0.58 occurring at

The flex boot remained attached to the outer boot ring 360 degrees

circumferentially, and showed no bonline separations. The cavity side of

the flex boot was evenly sooted and showed no evidence of flow or erosion

(Figures 62 through 64). It appeared typical of previous flight and static

test motor flex boots. A minimum of

circumference after motor burn. Table

affected depths and performance margins

PMS was 0.19 at 280 degrees.

3 NBR plies remained around the

7 presents the flex boot material

of safety (PHS). The worst case

STS-26A Fixed Housing

Overall views of the STS-26A fixed housing assembly are shown in Figures 6

and 7.

The fixed housing insulation erosion was smooth

wedgeouts of charred CCP material were observed

intermittently around the circumference

depth of these wedgeouts was 0.5 inch.

GCP.

and uniform. Post-burn

on the forward 2 inches

(Figure 65). The maximum radial

There was no heat effect to the

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There were no bondltne separations observed on the forward or aft end.

Photographs

Figures 66 through 69.

presented in Table 8.

stations. All margins

of the sectioned fixed housing assembly liner are presented in

Char and erosion analysis of the sections is

Figure 70 shows the location of the measurement

of safety were positive, with a minimum of 0.56


STS-26A Bearing Protector

The bearing protector was sooted along the entire length and circumference

(Figures 71 through 73). Slightly heavier soot and erosion was observed in

line with the cowl ring vent holes at the thickened portion, but there was

no bearing protector burn-through. There was no evidence of heat effect on

the Inner Diameter (ID) surface of the bearing protector.

STS-26A Flex Bearing

Exami,at|o. ot tile flex bearing revealed no damage, soot, heat effect, or

flow indications (see Figures 74 and 75). All rubber pads, metal shims,

and end rings appeared to be in nominal condition. Subsequent

refurbishment and testing has verified that the flex bearing is acceptable

for reuse.

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4.1.2 STS-26A Nozzle Internal Joints

Descriptions of the STS-26A nozzle internal joints follows.

STS-26A Aft Exit Cone-to-Forward Exit Cone (Joint No. I)

A cross-sectioned view of the STS-26A aft exit cone/forward exit cone field

joint is presented in Figure 76. Photographs of the post-flight joint are

shown in Figures 77 through 82.

The backfilled RTV extended below the joint char line circumferentially

except at the 266.2-degree location. RTV filled the radial ID portion of

the joint except at 236.2, 266.2, 292.8, and 296.6 degrees where unfilled

void areas were located. The backfill also reached the high pressure side

of the primary O-ring from 38 to 185, and 314.4 to 356.2 degrees. One

blowpath 0.I0 inch wide circumferentlally was observed at the 266.2 degrees

unfilled void area. The primary O-ring saw pressure, but showed no signs

of blowby, erosion, or heat effect.

Examination of the joint showed a black residue and aluminum oxide

corrosion appearing on both metal surfaces of the joint between the primary

and secondary O-rings, and outboard of the secondary O-ring intermittently

around the circumference. The black residue was heaviest from 131 to 270

to 0 degrees (Figure 83). The aluminum oxide corrosion was heaviest from 0

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to 90 to 131 degrees (Figure 8_).

been determined that the black

aluminum oxide corrosion.

There was no pitting observed. It has

residue is the beginning stage of the

The aft flange of the forward exit cone was scratched

location by a guide pin during aft exit cone demate

scratch was approximately 0.002 inch deep axially,

circumferentially and 0.375 inch wide radially.

at the 90-degree

(Figure 85). The

3.5 inches long

STS-26A Throat-to-Forward Exit Cone (Joint No. 4)

A cross-sectioned view of the STS-26A throat/forward exlt cone joint is

presented in Figure 86. Photographs of the post-flight joint are shown in


The RTV backfill extended below the joint char line and filled the axial

portion of the joint 360 degrees circumferentially. RTV reached the high

pressure side of the primary O-ring from 65 to 125, 195 to 210, and 312 to

350 degrees. One blow path measuring 1.0 inch circumferentially was found

at 310 degrees on the radial OD portion of the joint (Figure 93). The GCP

was sooted at this location, but not heat affected. The primary O-ring saw

pressure, but there was no evidence of blowby, erosion, or heat effect.

Soot was not evident on the radial ID or the axial portions of the joint.

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It is believed that the blow path extended cohesively through the RTV at

this location. White deposits, possibly salt, were found on the phenolic

radial ID portion of the joint intermittently around the circumference.

Corrosion was evident on both metal surfaces

25 to 125, and 190 to 330 degrees on

circumferentially on the forward exit cone.

on the metal surfaces.

of the joint, extending from

the throat and 360 degrees

There was no pitting observed

STS-26A Nose Inlet-to-Throat (Joint No. 3)

A cross-sectioned view of the STS-26A nose inlet/throat joint is presented

in Figure 94. Photographs of the post-flight joint are shown in Figures 95

through lO0.

The RTV backfill extended below the joint char line 360 degrees

circumferentially. RTV completely filled the radial ID portion of the

joint circumferentially except from 309 to 313 degrees. RTV also extended

onto the radial OD from 52 to 70, 146 to 149, 163 to 171, 174 to 190, 195

to 197, and 210 to 220 degrees, but did not reach the primary O-ring. One

blow path measuring 0.9 inch wide circumferentially and 1.2 inches deep

radially was observed at 136 degrees (Figures 101 and 102). The blow path

terminated within the RTV. The primary O-ring did not see pressure.

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Aluminum oxide corrosion was observed on both metal surfaces of the joint

inboard of the primary O-ring, but no pitting was observed. Rust was found

within the metal/adhesive separations on the forward end of the throat

support housing intermittently around the circumference.

Minor surface scratches were observed on the aft end of the nose inlet

housing (-504 ring aft end) where jacking screws were used to disassemble

the joint.

STS-26A Nose Inlet-to-Bearing Forward End Ring-to-Cowl (Joint No. 2)

A cross-sectioned view of the STS-26A nose inlet/forward end ring/cowl



The RTV extended below the joint char line and filled the axial portion of

the joint 360 degrees clrcumferentlally.

nose cap and cowl showed RTV mixed with

circumferentially. The adhesive was

The radial bondline between the

the EA913 NA adhesive 360 degrees

typically sandwiched between two

layers of RTV. There was no RTV extending to the primary O-ring. One blow

path was observed at 216 degrees. On the aft end of the nose cap, the blow

path measured 0.60 inch wide circumferentiaily and charred the GCP

approximately 0.005 inch deep axially. On the forward end of the cowl

ring, the blow path measured 0.40 inch wide circumferentially and charred

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the sllica cloth phenollc (SCP) approximately 0.01 inch deep axially

(Figure 113). The EA913 NA adhesive on the cowl eroded approximately 0.1

inch deep axially (maximum) by 0.7 inch vide circumferentially at the blow

path location. Soot was observed on the nose cap/forward end ring

interface surfaces, reaching the primary O-ring from 156 to 162, and 180 to

240 degrees (Figure 112). There was no blowby, erosion, or heat effect to

the primary O-rlng. Soot also extended to midway between the bolt holes

around the remainder of the circumference.

Both the nose inlet housing and the cowl housing metal surfaces were

heavily sooted at the blow path location. Electrical conductivity tests

run on these parts showed that there was no heat damage. The bearing

forward end ring was also sooted, and the paint was chipped off in various

spots, but neither the end ring or the paint were heat affected.

Water was found on the nose housing aft face and in the bolt holes from 12

to 198 degrees. Aluminum oxide corrosion was observed on the forward face

of the cowl housing from 214 to 224 degrees and extended approximately 0.5

inch radially inward. Corrosion and salt deposits were also found on the

ID surface of the cowl housing

circumferentlally. This indicates water

and bearing protector during splashdown.

forward flange 360 degrees

leaked between the cowl housing

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STS-26A Fixed Housing-to-Bearing Aft End Ring (Joint No. 5)

A cross-sectloned view of the STS-26A aft end ring/fixed housing joint is



RTV filled approximately 80 percent of the axial portion of the joint and

reached the high pressure side of the primary O-ring at 25 to 30, 35 to 43,

55, 65 to 78, 240, and 308 to 313 degrees. Voids isolated within the RTV

were observed on the radial portion of the joint intermittently around the

circumference (Figure 121). The largest measured 0.9 inch deep radially by

1.7 inches wide circumferentially. None of the voids extended to the flex

boot cavity. There were no blow paths observed in the joint.

Water was found on the aft face of the aft end ring and in the bolt holes

intermittently around the circumference. Rust corrosion was observed on

both metal surfaces of the joint between the O-rings at 15 degrees (Figure

122), but there was no pitting. Rust corrosion was found on the aft end

ring inboard of the secondary O-ring intermittently around the

circumference. Again, no pitting was observed. A white corrosion spot

(0.10 inch in diameter) located at 260 degrees was also noted (Figure 123).

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4.2 STS-26B Nozzle/Flex Bearing

Overall erosion of the STS-26B forward nozzle assembly CCP ablative liner

was smooth and uniform. All CCP delamlnations, wedgeouts, and pop-ups were

determined to be post-burn occurrences resulting from cooldown of the

liners. Blovpaths were observed in joints 2 and 4, but there was no

blowby, erosion, or heat effect to the primary O-rings. Small amounts of

corrosion were found on the metal surfaces of joints 1, 2, 3, and 4, but no

pitting was observed. Beavy corrosion and pitting was found on the nose

inlet housing bonding surfaces when the phenolics were washed off. The

forward exit cone also showed corrosion on the ID bonding surface.

Post-flight subassembly flow surface gaps are shown in Figure 124. Overall

views of the nozzle are shown in Figures 125 through 130.

4.2.1STS-26B Nozzle Components

STS-26B Aft Exit Cone Assembly

An overall view of the STS-26B aft exit cone fragment is shown in Figure

131.

The aft exit cone was severed aft of the compliance ring by the LSC. The

nozzle severance system performance was nominal. The exit cone cut was

clean, with no unusual tearing or breaking. The remaining aft exit cone

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fragment showed missing CCP liner 360 degrees circumferentially. This is a

typical post-fllght observation and occurs at LSC firing and during

splashdown. GCP plles exposed by the missing liner showed no signs of heat

effect.

The polysulfide groove fill on the forward end of the aft exlt cone showed

one separation between the polysulfide and the GCP Insulator. The

separation was located at 211 degrees and measured 0.02 Inch wide radially,

0.04 inch deep axlally and 1.3 inches long circumferentlally. Post-flight

measurements of the polysulfide groove radial width (Table 9) show that the

GCP insulator did not pull away from the aluminum shell during cooldown.

The polysulfide shrank axially aft up to 0.I0 inch.

There were no separations observed within the GCP insulator on the forward

end.

STS-26B Forward Exit Cone Assembly

Overall views of the STS-26B forward exit cone ale shown in Figures 129 and

130.

The forward exit cone showed missing CCP liner over the center 14 inches of

the cone 360 degrees circumferentially. This is a typical post-flight

observation and occurs at splashdown and during Diver Operated Plug (DOP)

insertion. The GCP insulator exposed by the missing liner showed no signs

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of heat effect. The CCP liner remained bonded on the forward 13 inches and

on the aft 9 inches of the cone. These portions showed nominal erosion

with no major washing or pocketing. The aft 9 inches of the liner showed

the typical dimpled erosion pattern that has occurred on all flight and

static test forward exit cones (Figure 132). The maximum radial depth of

the dimpled erosion was 0.15 inch.

The aft end of the forward exit cone showed no bondline or cohesive

_eparations. Bondline separations on the forward end of the forward exit

cone were noted between the steel shell and the EA946 adhesive

circumferentially except at 105 degrees. Separations were also found

between the GCP and CCP, within the GCP, and within the adhesive. Figure

133 lists the location and radial width measurements of all separations on

the forward exit cone. These separations are typical observations seen on

previous static test and flight nozzles and have been shown to occur

post-burn.

Photographs of the sectioned forward exit cone liner are presented in




occurring at station 1 (0 and 180 degrees), and station 8 (270 degrees).

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Following washout of the pheno]ics, large areas of corrosion were noted

along the forward 5 to 12 inches of the ID bonding surface (Figures 138 and

139). This "band" of corrosion appeared aft of the forward shear pins.

hight and dark colored areas of corrosion as well as rust spots and pitting

were observed. Corroded areas were also found on the aft 7 inches of the

ID bonding surface centered around the aft shear pin holes (Figures 140

through 144). The largest area was at 120 degrees (Figure 142). Visual

inspections of these indicate that sea water may have leaked through the

shear pin holes where the lightning cables were attached (every 30

degrees). Light and dark areas of corrosion, rust spots, and pitting were

observed. Small rust spots were also noted intermittently around the rest

of the ID surface (Figure 142). These were typically 0.050 to 0.10 inches

in diameter. This hardware will be inspected during refurbishment for

compliance to STWT-3434 (Refurbishment Of And Acceptance Criteria For Space

Shuttle SRM Nozzle Metal Hardware).

STS-26B Throat Assembly

Overall views of the STS-26B throat assembly (throat ring and throat inlet

ring) are shown in Figures 125 and 126.

The throat post-flight mean diameter was 55.876 inches (erosion late ot

8.18 mils/second based on an action time of 123.2 seconds). Nozzle

post-burn throat diameters have ranged from 55.787 to 56.38 inches. The

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flow surface bondline gap between the throat and throat Inlet

0.10 inch and is typical of past static test and flight nozzles.

rinEs was

The throat and throat inlet rings eroded smoothly with no pockets or major

washes observed. The forward end of the throat inlet ring showed post-burn

wedgeouts of charred CCP material from 28 to 40, 95 to 105 and 355 to 0 to

5 degrees (Figure 145). The maximum axial width of the vedgeouts was 0.75

inch at the 28-to-40-degree location. Post-burn wedgeouts of the throat

inlet ring forward end have been observed on previous post-fllght nozzles.

The forward 1.5 inches of the throat ring showed popped-up charred CCP

material intermittently around the circumference. Sharp edges indicate the

popped-up material occurred after motor operation. Marks resulting from

DOP insertion were observed on the throat ring intermittently around the

circumference.

Bondline separations on the aft end of the throat ring between the EAgI3 NA

adhesive and the steel throat support housing were observed around the

majority of the circumference. Separations were also found between the

adhesive and GCP, within the GCP, and within the adhesive. There were no

separations between the GCP and CCP on the aft end.

throat inlet ring showed metal to adhesive

circumferentially except from 255 to 0 degrees.

observed between the GCP and CCP. Figure 146 lists the location and radial

width measurements of all separations on the throat assembly. These

The forward end of the

bondline separations

Separations were also

.,v,s,o. __ ooc .o. TWR- 17 2 7 2 ] vo.

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separations are typical observations seen on previous

flight nozzles and have been shown to occur post-burn.

static test and

Photographs of the sectioned throat assembly liner are presented in Figuze_

147 through 149. Char and erosion analysis of the sections is presented in

Table 11. Figure 25 shows the location of the measurement stations. All

margins of safety were positive, with a minimum of 0.07 occurring at


STS-26B Nose Inlet Assembly

Overall views of the STS-26B nose inlet assembly (forward nose ring, aft

inlet ring, and nose cap) are shown in Figures 125 through 128.

The ply angle of the forward nose ring was checked and found to be of the

RSRM design. The flow surface bondline gap between the forward nose (-503)

ring and the aft inlet (-504) ring was 0.18 inch. The flow surface bondline

gap between the -503 ring and nose cap was 0.05 inch. These post-flred

measurements are typical of past static test and flight nozzles.

The -503 and -504 rings showed smooth erosion with no pockets or major

washes observed. The -503 ring showed popped-up charred CCP material at

the nose cap interface from 155 to 165 degrees. The popped-up material was

0.08 inch wide axially and occurred ariel motor operation. Impact mark:;

TWR-17272!

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occurring after motor operation were observed on both rings inter_Jttently

around the circumference (Figures 150 and 151). The marks most likely

resulted from the loose aft and forward exit cone CCP material in the motor

at splashdown.

The nose

observed.

post-burn

cap showed smooth erosion with no pockets or major washes

The aft 2.0 to 3.5 inches of the nose cap showed typical

wedgeouts intermittently around the circumference (Figure 152).

These measured approximately 0.5 in. deep radially at the cowl interface.

One wedgeout location from 5 to 20 degrees showed slag covering exposed CCP

material. Sectioning of the liner determined that this wedgeout occurred

post-burn.

The aft end of the nose inlet assembly (-504 ring aft end) shoved metal to

adhesive bondline separations measuring 0.02 inch wide radially from 238 to

245 degrees, and at 250 degrees. There were no cohesive separations or

separations at the adhesive/GCP and GCP/CCP interfaces. Bondline

separations were observed on the aft end of the nose cap between the metal

and EA946 adhesive at 105, 135 to 255, 285 to 315, and 345 degrees. These

separations were typically 0.005 inch wide radially. There were no

cohesive separations or separations at the adhesive/GCP and GCP/CCP

interfaces. Figure 153 lists the location and radial width measurements of

all separations on the nose inlet assembly. These separations are typical

oh_ervatlons seen on previous static test and f]igh¢ nozzle_, and have been

shown to occur post-burn.

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Photographs of the sectioned nose inlet assembly rings are presented in

Figures 154 through 161. Char and erosion analysts of the sections is

presented in Tables 12 and 13. Figure 38 shows the location of the


of O.O1 occurring at station 32 (180 degrees) for the -503/-504 rings, and

0.01 occurring at station 24 (225 degrees) for the nose cap.4

L_

Following the washout of the phenollcs, it was found that the aluminum nose

inlet housing had extensive corrosion and pitting on all bonding surfaces

360 degrees circumferentlally (Figures 162 through 167). Most of the

corrosion on the nose cap bonding surface was found on the aft 5 inches 360

degrees clrcumferentially. The forward edge of this corrosion was shaped

in a "saw tooth" pattern (Figure 166). The leading edge of the nose inlet

housing showed areas of pitting approximately 0.04 to 0.05 inch deep

(Figure 168). The entire -503 ring bonding surface was heavily corroded

and pitted 360 degrees clrcumferentially, and approximately 90 percent of

the -504 ring bonding surface showed various stages of corrosion and

pitting. This corrosion has been attributed to seawater which enters

bondline separations during splashdown and retrieval (Ref. Memo

L231-FY89-MI30). The metal bonding surfaces were not accessible until

phenolic washout at Clearfield Operations. Therefore, corrosion protection

was not applied to these surfaces until approximately 4 months after

flight. This hardware will be inspected during refurbishment for

compliance to STW7-3434 (Refurbishment Of And Acceptance Criteria For Space

Shuttle SRM Nozzle Metal Hardware).

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STS-26B Cowl Ring

Overall views of the STS-26B cowl ring are shown in Figures 127 and 128.

Close-up views are shown in Figures 169 through 171. All cowl vent holes

appeared plugged with slag on the OD of the ring (Figure 172).

The cowl ring showed typical ridged erosion intermittently around the part

circumference. The forward portion of the ring eroded a maximum of 0.15

inch greater than the aft portion (Figure 169). This is a result of the

low ply angle of the cowl ring and has been observed on the majority of

flight and static test nozzles. One wedgeout was observed on the aft 3.5

inches of the cowl ring from 120 to 137 degrees (Figure 173). The maximum

radial depth of the wedgeout was 0.6 inch at the outer boot ring interface.

Slag coated the exposed CCP material at the wedgeout location. Sectioning

of the liner determined that this wedgeout occurred post-burn.

There were no bondline separations on the forward end of the cowl ring.

Photographs of the sectioned cowl ring are presented in Figures 174 through

177. Typical subsurface ply lifting was observed intermittently around the

circumference along

separation was 0.20

evidence of flow or

analysis

the length of the cowl. The largest ply lift

inch at 270 degrees (Figure 177). There was no

erosion within the delaminations. Char and erosion

of the sections is presented in Table 14 (Stations 0 through 7).

REVISION _CN°TVR-17272 I _L

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Figure 55 shows the location of the measurement stations. All margins of

safety were positive, with a minimum of 0.04 occurring at station 0 (45

degrees).

STS-26B Outer Boot Ring/Flex Boot

Overall views of the STS-26B outer boot ring are shown in Figures 127 and

128. Close-up views are shown in Figures 169 through 171. The bondline

between the outer boot ring and cowl ring remained intact with no

indications of flow. The flow surface bondline gap was 0.20 inch and is

typical of past static test and flight nozzles.

The structural backup outer boot ring was intact. The flow surfaces showed

smooth erosion with no pockets, wedgeouts, or major washes. Minor wash

area_ extended from the cowl to the forward 1.5 inches of the OBR from 120

to 150, and 151 to 158 degrees, and measu[ed a maximum o_ 0.2 inch radially

deep. These have occurred on the majority of flight and static test

nozzles. Popped-up charred CCP material was observed on the forward 1.8

inches of the OBR intermittently around the circumference. The popped-up

material is a common observation and occurs after motor operation.

Delaminations in the charred CCP of the aft tip were observed 360 degrees

circumferentially (Figure 178). Charred CCP material on the aft tip

adjacent to the flex boot fractured and popped up over a majority of the

circumference.

TWR-17272 I

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Photographs of the sectioned outer boot ring are presented in Figures 179

through 182. Char and erosion analysis of the sections is presented in

TabIe 14 (Stations 8 through 11.3). Figure 55 shows the location of the


of 0.59 occurring at station 10 (0 degrees).

The cavity side of the flex boot was evenly sooted and showed no evidence

of flow or erosion (Figures 183 through 185). It appeared typical of

previous flight and static test motor flex boots. A minimum of 3.0 NBR

plies remained around the circumference after motor burn. Table 15

presents the flex boot material affected depths and Performance Margins of

Safety. The worst case PMS was 0.19 at 280 degrees.

STS-26B Fixed Housing Assembly

Overall views of the STS-26B fixed housing assembly are shown in Figures

127 and 128.

The fixed housing insulation showed smooth erosion with no pockets or major

washing observed. Post-burn wedgeouts of charred CCP material were

observed on the forward 2.0 inches of the fixed housing insulation from 30

to 65, 135 to 145, and 165 to 180 degrees. The wedgeouts were a maximum of

0.5 inch deep radially. There was no heat effect to the GCP.

TUR- 17272 IREVISION DOC NO VOLI

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There were no bondltne separations observed on the forward or aft end.

Photographs of the sectioned fixed housing assembly liner are presented in





STS-26B Bearing Protector

The bearing protector was sooted along the entire length and circumference

(Figures 190 through 192). Heavier soot and erosion were observed in line

with the cowl ring vent holes at the thickened portion of the bearing

protector. Erosion depths at the vent hole locations are presented in

Table 17. There was no evidence of heat effect on the IO surface of the

bearing protector.

STS-26B Flex Bearing

Examination of the flex bearing revealed no damage, soot, heat effect, or

flow indications (Figure 193). All rubber pads, metal shims, and end rings

appeared to be in nominal condition. Subsequent rerfurbishment and testing

has verified that the £1ex bearing is acceptable [o[ reuse.

.,v,s,o.__ ooc.o TVR- 17272 IvoL

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SPACE OPERATIONS

4.2.2 STS-26B Nozzle Internal Joints

Descriptions of the STS-26B nozzle internal joints follows.

STS-26B Aft Exit Cone-to-Forward Exit Cone (Joint No. 1)

A cross-sectioned view of the STS-26B aft exit cone-to-forward exit cone

field joint is presented in Figure 194. Photographs of the post-flight

joint are shown in Figures 195 through 200.

The backfilled RTV extended below the joint char line 360 degrees

circumferentlally. RTV filled the radial ID portion of the joint except at

103 degrees where an unfilled void area approximately 1.0 inch wide

circumferentially was located. The backfill also extended to the high

pressure side of the primary O-ring from 0 to 81, 82 to i01, 103 to 123,

154 to 178, 182 to 237, 243 to 251, 258 to 265, and 268 to 0 degrees.

There were no blowpaths observed in the joint and the primary O-ring saw no

pressure. Char was observed on the RTV in the axial portion of the joint

at 237 degrees. The RTV was not eroded or heat affected at the charred

location. It is believed that the char penetrated the joint at splashdown.

Examination of the joint showed a black residue and aluminum oxide

corrosion appearing on both metal surfaces between the primary and

secondary O-rings, and outboard of the secondary O-ring intermittently

,,v,s,on__ oocNO TWR- 17272 IvoLSEC 1"PAGE

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around the circumference (Figure 201).

heaviest from 112.6 to 143.2 degrees. There

was determined that the black residue is

aluminum oxide corrosion.

The aluminum oxide corrosion was

was no pitting observed. It

the beginning stage of the

One through hole on the forward exit cone housing aft flange was dinged by

a guide pin during the aft exit cone demate. The ding was approximately

0.02 inch deep and was located at 95.6 degrees (Figure 202).

STS-26B Throat-to-Forward Exit Cone (Joint No. 4)

A cross-sectioned view of the STS-26B throat-to-forward exit cone joint is



The RTV backfill extended below the joint char line and filled the radial

ID portion of the joint circumferentially, except at 185 degrees. RTV

filled the axial portion of the bondline from 40 to 165 degrees, and 240 to

345 degrees. RTV did not reach the high-pressure side of the primary

O-ring.

185 deg.

resulting

ther was no evidence of blovby, erosion,

showed no signs of heat effect.

One blow path measuring 0.25 inch circumferentially was found at

Excess grease at this location inhibited the RTV backfill,

in an unfilled void area. The primary O-ring saw pressure, but

or heat effect. The GCP also

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Rust corrosion was observed on both surfaces of the joint within the metal

housing�adhesive bondltne separations intermittently around the

circumference. Black corrosion was observed near the primary sealing

surface of the throat housing aft end from 80 to 85 degrees, and 345 to 0

to 4 degrees. There was no pitting on the metal surfaces.

STS-26B Nose Inlet-to-Throat (Joint No. 3)

A cross-sectioned view of the STS-26B nose inlet-to-throat joint is



The RTV backfill extended below the joint char line 360 degrees

circumferentially. RTV filled the radial ID portion of the joint

circumferentially except at 50 degrees. RTV also extended onto the radial

OD up to the GCP/CCP interface at 35, 275, and 325 degrees. An unfilled

void area, 1.0 inch circumferentially, was located at 50 degrees. There

was no blow path to the void area. The primary O-ring did not see

pressure. Grease was observed on both sides of the joint 360 degrees

circumferentially extending 0.I to 1.0 inch inboard of the primary O-ring.

Minor surface corrosion was observed on the aft end of the nose inlet

housing inboard of the primary O-ring, but no pitting was observed. This

aluminum oxide corrosion extended approximately half way down the ID side

TWR-17272I

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Of the primary O-ring groove at 325 degrees. There was no corrosion

on the forward end of the throat housing.

STS-26B Nose Inlet-to-Bearing Forward End Ring-to-Cowl (Joint No. 2)

A cross-sectloned view of the STS-26B nose inlet-to-bearing forward end

ring-to-cowl joint is presented in Figure 217. Photographs of the

post-flight joint are shown in Figures 218 through 226.

The RTV extended below the joint char line and filled the axial portion of

the joint 360 degrees circumferentially. The radial bondline between the

nose cap and cowl showed RTV mixed with the EA913 NA adhesive

intermittently around the circumference, The adhesive was typically

sandwiched between two layers of RTV. RTV filled approximately 80 percent

of the axial bondline between the nose cap and bearing forward end ring.

No RTV extended to the primary O-ring. One blow path was observed at 266

degrees and measured 0.5 inch wide circumferentially (Figure 227). The

cowl SCP and nose cap GCP insulators showed no heat effect. The primary

O-ring saw pressure, but there was no evidence of blowby, erosion, or heat

effect. Soot was observed on the radial OD of the joint 360 degrees

circumferentially. Soot reached up to the axial bolt holes on the nose

inlet housing intermittently around the circumference, but did not reach

the primary O-ring. Salt deposits were also noted on the radial OD

surfaces.

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Both the aft face of the forward end ring flange and the forward face of

the cowl housing were sooted at 130 to 153, 165, 255, and 303 to 310

degrees, but were not heat effected. The paint on the forward end ring

OD flange surface was chipped off in various spots, but was not heat

affected. Minor rust spots were noted in areas where the paint was chipped

off. Aluminum oxide corrosion was observed on the forward end ring/nose

inlet housing interface surfaces, and on the cowl housing/forward end ring

interface surfaces intermittently around the circumference. Aluminum oxide

corrosion was also found intermittently on the forward flange ID surface of

the cowl housing (Figure 228).

STS-26B Fixed Housing-to-Bearing Aft End Ring (Joint No. 5)

A cross-sectioned view of the STS-26B aft exit cone/forward exit cone field



RTV filled approximately 75 percent of the axial portion of the joint and

reached the high pressure side of the primary O-ring from 75 to II0, 123 to

128, 135 to 148, and 195 to 218 degrees. Voids isolated within the RTV

were observed on the radial portion of the joint intermittently around the

circumference. The largest measured 0.45 inch deep radially by 0.30 inch

wide circumferentially. A void at 171 degrees extended onto the axial

portion of the joint, but terminated within the RTV. There were no blow

paths observed in the joint, and the primary O-ring did not see pressure.

.EV,S,ON-- OOCNO.TVR- 17 2 7 2 [ vol

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Rust corrosion was found on the aft end ring inboard of the secondary

O-ring intermittently around the circumference. No pitting was observed.

4.3 Instrumentation

There was no instrumentation installed on the STS-26 nozzles.

5.0 DISCREPANCY REPORTS AND PROCESS DEPARTURES

The STS-26 Nozzle DRs and PDs reviewed by the Morton Thiokol senior

material review board are included in Appendix A. These were presented in

the STS-26 RSRM Acceptance

descriptions of the DRs and

observations are discussed below.

Review Level IIl (TWR-18117A). Brief

PDs, and correlations to post-flight

5.1STS-26A DRs and PDs

Aft Exit Cone

DR 123524-01 (Waiver No. RW'_/404)

LDIs within the GCP were found at 240 degrees, 54 inches aft of the forward

end. This portion of the aft exit cone was severed by the LSC during

reentry and wa_ not recovered. Post-flight inspection of thi_ parr wa_ not

possible.

TWR-17272 I

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SPACE OPERATIONS

DR 162635-01, -02 (Waiver No. RW-g 405)

LDIs within the GCP were found at 222 degrees (39.5 inches aft of the

compliance ring), and 240 and 243 degrees (4 inches aft of the compliance

ring). This portion of the aft exit cone was severed by the LSC during

reentry and was not recovered. Post-flight inspection of this part was not

possible.

Forward Exit Cone

DR 151717-O1 (Waiver No. RWW 387)

Eight LDIs within the GCP were found running 360 degrees circumferentially

along a full ply length. Post-fllght inspection of the exposed GCP did not

reveal any delaminations extending to the surface.

PD 150663-01

The white stripe (90-degree mark) on the phenolic liner was 1.75 inches

from the 90-degree reference pin (0.75 inch over maximum). The liner was

bonded at the same radial location as the dry fit. Orientation to

correlate post-flight performance with any pre-flight anomalies was not

affected.

Throat Assembly

RE VISION

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DR 128578-01

Intermittent pitting, a maximum of 0.002 inch deep, was found on the aft

sealing surface. After being repaired, the joint was succesfully leak

tested. Post-flight inspection did not reveal any indications of blow-by.

Nose Inlet Assembly

DR 152142-01

Phosphoric Acid Anodization (PAA) and EA9228 Primer applied to the bonding

surfaces was not uniform (dark streaks and spots). The PAA and primer

system was deleted from the engineering design change. This part was grit

blasted and the phenolics bonded using 51-L surface preparation techniques.

All of the phenolics were intact and remained bonded to the housing.

PD 150024-01

The EA946 adhesive for the nose cap bond was not applied within 6 hours

from the grit blast requirement. This was reworked to blueprint

requirements. No post-flight anomamlies were noted.

PD 150024-02

The EA946 adhesive for the nose cap bond was not applied within 1 hour from

methylchloroform wipe. This was reworked to blueprint requirements. No

post-flight anomamlies were noted.

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Cowl Assembly

DR 126842-01

Intermittent pitting was found on the cowl housing (worst-case condition

was 0.180 inch in diameter by 0.039 inch deep on the ID flange. The pits

were honed out to remove sharp edges. No post-flight anomalies were noted.

Flex Bearing

DR 123208-01

One threaded hole (0.190-32 UNF) on the aft end ring accepted the no-go

threaded plug gage for 6.5 turns. Proper bolt torque was verified and

showed no damage. Post-flight inspection showed damage to the bolt hole.

DR 123439-01

The unbond area on pad ii exceeded the maximum

flex bearing passed all of the acceptance

refurbishment requirements.

allowable of 9 in2. The

test_ and post-flight

5.2 STS-26B DRs and PDs

Nozzle/A[t Segment Assembly

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DR 153960-01

A broken girth gage wire was found between the aft dome boss and nozzle

assembly. This did not affect O-ring gap openings. The joint successfully

passed leak check. Post-flight inspection showed no anomalies as a result

of this condition.

Aft Exit Cone

DR 123533-01, (Waiver No. RWW 406)

An LDI measuring 2.35 inches circumferentially, 1.2 inches axially, and

0.031 inch radially was found in the GCP 0.393 inch from the forward end at

45 degrees. Post-fllght inspection of the GCP after sectioning (Figure

236) did not reveal any delaminatlons.

Throat/Nose Inlet Joint Assembly

DR 150682-01, -02

The primary to secondary seal cavity was pressurized to

high pressure leak check (the requirement is 740 ± 15 psig),

during low pressure leak check (the requirement is 30 ± 3 psig).

passed leak check, and no anomalies were observed during

inspection.

1040 psig during

and 40 psig

The joint

post-flight

.EV,S,ON OOCNO.TWR- 17 2 7 2I VOL

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Nose Inlet Assembly

PD 149145-01

This involved the forward first wrap of the nose ring during the carbon

hydroclave cure. While decreasing pressure, the pressure dropped to 168

psig for 4 minutes, then remained in tolerance during the remainder of the

cure. The CCP liner met all acceptance criteria. There were no anomalies

observed during post-flight inspection.

Cowl Assembly

DR 128474-01

Pitting was observed on the OD and ID surfaces of the cowl housing.

Maximum depths were 0.049 inch on the ID side, and 0.041 inch on the OD

side. These were blended out to remove sharp edges. Post-flight

inspection has not revealed any anomalies.

Flex Bearing

DR 123437-01

One threaded hole (0.750-16 UNF) on the aft end ring accepted the no-go

threaded plug gage for eight turns. A helical coil insert was used as a

standard repair. There was no damage noted during post-flight inspection.

The repair did not affect the flex bearing performance.

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Bearing Protector

PD 127767-01

This involved the GCP autoclave cure for the bearing protector inner ring.

The autoclave vacuum dropped below the minimum of 15 in. Hg for a total of

177 minutes. The inner ring met all acceptance tests, and the bearing

protector assembly using this ring passed strength tests. There were no

anomalies observed during post-flight inspection.

6.0 NOZZLE

PROGRAM REVIEW BOARD (RPRB) ASSESSMENT

The NCPT reviewed all observations documented in this report. The

classified five Problem Reports (written at KSC) as minor anomalies.

internal nozzle joint inspections at Clearfield,

classified five observations as potential anomalies.

further classified as minor anomalies, and the

COMPONENT PROGRAM TEAM (NCPT) RECOMMENDATIONS AND REDESIGN

team

After

the team initially

Three of these were

other two remained

observations. These were presented to the RPRB on 9 and II November, 1988.

The RPRB agreed with all the classifications. These minor anomalies were

recorded on Post-Fire Anomaly Record (PFAR) forms and are included in

Appendix B. The PFARs contain detailed descriptions and corrective actions

as accepted and/or modified by the RPRB. A listing of the PFARs is listed

below.

6.1STS-26A Nozzle

PFAR NUMBER

360L001A-11

DESCRIPTION

Corrosion on aft exit cone metal between primary

and secondary O-rings.

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SPACE OPERATIONS

360L001A-12

360L00IA-43

6.2 STS-26B Nozzle

PFAR NUMBER

360L00IB-IO

360L00IB-38

360LOOIB-42

360LOOIB-44

360LOOIB-45

Corrosion on forward exit cone metal between

primary and secondary aft exit cone O-rings.

RTV and EA913 NA adhesive mixing in joint 2.

DESCRIPTION

Corrosion on aft exit cone metal between primary

and secondary O-rings.

Ding on forward exit cone aft flange.

Corrosion on forward exit cone metal between

primary and secondary aft exit cone O-rings.

RTV backfill in joint 4 inhibited by excessive

grease.

RTV and EA913 NA adhesive mixing in joint 2.

REVISION oocNo TUR-17272 I VO[

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MORTON THIOKOL INC

Space Operations

Flex Bearing

Support Housing

Throat Inlet Housing

Throat Ring

Throat Inlet Ring

-504 Aft InletRing

-503Nose Ring

_ft End Ring

Snubber Assembly

Forward Exit Cone Liner

Forward ExitCone Housing

Nose InletHousing

FixedHousing

-402 Nose Cap Fixed Housing

Forward Insulation

End Ring Cowl

Flexible Boot Ring

Structural Backup OBR

Bearing Protector

Cowl Housing

r- Aft Exit Cone Jner

(carbon-cloth phenolic)

/ /-- Aft Exit Cone Shell

5-T73 aluminum)Q

r,,-_ 21_ _-,__ r- Aft Exit Cone'v "'" --___._ II_J_A__ / Structure

Actuator Bracket \ _-kin_r--C:;haoed Charae __(7075-T73 aluminum

• Compliance Ring(7075-T73 aluminum) A017773a

Figure 1 STS-26 Nozzle Components

oocNo TWR-17272 Ivo_

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Space Operations

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Space Operations

f O.20 In.

0.22 in. "-I

_ --.....

STS-26A Joint Flow Surface Gap Openings

OOCNOTWR-17272

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Figure 10 STS-26A Aft Exit Cone Fragment

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AngularLocation

C153045607590

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30 0.015 Metal/Adhesive45 0.015 Metal/Adhesive60 0.002 Metal/Adhesive

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Figure 13 STS-26A Forward Exit Cone Bondline Separations

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Table 4 STS-26A Nose Inlet Rings (-503, -504) Rros_on and Char Data

Angular Location

0 degrees

Measured Erosion

Measured Char

Adjusted Char t

2E + 1.25AC

RSRM Kin Liner Thkns

Margin of Safety

Stations

28 30 32 34 36 38 39.5

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2.55 2.44 2.26 2.36 2.48 2.82

3.252 2.g50 3.182 3.200 3.026 2.981

0.28 0.21 0.41 0.35 0.22 0.06

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Measured Char

Adjusted Char*

2E + 1.25AC

RSRM MAn Liser Thkns

Margin of Safety

1.09 0.83

0.70 0.70

0.53 0.53

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0.41 0.41 0.23 O.10

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Measured Char

Adjusted Char*

2E + 1.25AC

RSRM Min Liner Thkns

Margin of Safety

0.96 0.75 0.86 0.82

0.85 0,75 0.76 0.69

0.64 0.56 0.57 0.52

2.72 2.20 2.43 2.29

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0 67

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2 29

3 200

0 40

0.90 1.20

0.70 0.48

0.53 0.36

2.46 2.85

3.026 2.981

0.23 0.05

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Measured Char

Adjusted Char*

2E + 1.25AC

RSRM Min Liner Thkns

Margin of Safety

1.08 0.87 0.94

0.88 0.72 0.67

0.66 0.54 0.50

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0 69

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2 37

3 200

0 35

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0.62 NA

0.47 NA

2.46 NA

3.026 2.981

0.23 NA

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Measured Erosion 1.15 0.87 0.89 0.82 0.86 0.91 0.97

Measured Char 0.78 0.62 0.78 0.62 0.60 0.67 0.65

Adjusted Char" 0.59 0.47 0.59 0.47 0.45 0.50 0.49

2E + [.25AC 3.03 2.32 2.51 2-22 2.28 2.45 2.55

RSRM Min Liner Thkns 3.508 3.252 2.950 3.182 3.200 3.026 2.981

Margin of Safety 0.16 0.40 0.17 0.43 0.40 0.24 0.17

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Table 12 STS-26B Nose Inlet Rings (-503,-504) Erosion and Char Data

Angular Location

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Stations

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2E + 1.25AC 2.83 2.34 2.44 2.20 2.34 2.52 2.54

RSRM Min Liner Thickness 3.508 3.252 2.950 3.182 3.200 3.026 2.981

Mscgin of Safety 0.24 0.39 0.21 0.45 0.37 0.20 0.17

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0.36 0.20 0.35 0.30 0.19 NA

180 degrees

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Measured Char 0.77

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Degree Remaining Affected Depth MarginLocation Plies @in.) Of Safety*

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STS-26B Maximum Bearing Protector Erosion in Line with Covl Vent F[oles0.2400.2300.2200.2100.2000.1900.180

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0 0.08110 0.04020 0.05630 0.04540 0,07750 0,06960 0.13670 0.08180 0.15990 0.237

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OF POOR QUALITY

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TWR-17272

A-40

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MORTON THIOKOL INC

Space Operations

Appendix B

REVISION

TWR-17272OOC NO.

SECI PAGE

I VOL

B-1

POSTFIRE ANOMALY RECORD (PFAR)

...................... . ............ . .... . .... . .... .°.. ......................... . ...................... . ...................

1. PFAR NUMBER I 3. INSPECTION LOCATION 4. REFERENCE SQUAWK NUMBER 5. REFERENCE PR _UMBER

36DLOO1A-11 t KSC X T-24/T-97 -.26"0045 . PV6-111293

........ °... .... • ......... . ..... . ... .......... ° ........................ , .... , .....

2 sRM,OTOR,UMBER ,7 ,2 6 IF,.O.BER 7 REFERENCESPRNUMBER36OLOO A ...... N/A f N/A

..................... . ............ . .... . .... . .... . ........................................... . ..........................

8. TITLECORROSION BETWEEN PRIMARY AND SECONOARY AFT EXIT CONE O-RINGS

........ o ................. ° ......................... . .............................. . ....................................

9. CLASSIFICATION

OBSERVATION MINOR ANOMALY X MAJOR ANOMALY CRITICAL ANOMALY.... , .......

........ . ....... . .... . ............ o .... . ......... °... ................. . .................................................

10. PART NUMBER I 11. SERIAL NUMBER I 12. PART DESCRIPTION

1U76039"01 I 0000001 I AFT EXIT CONE ASSY........ .°.. .... ° .... . ................. . ................. . .................................... .o. ............ • ..........

13. REPORTED BY (NAME / ORGANI_TION / OBSERVATION DATE)

K. B. NIELSEN / MTI QA / 10107/88................ ° ...................... . ............ ° ............ • .... ° ....... ° .... =... ................... ° .............

14. RESPONSIBLE C_PONENT TEAM / PROGRAM MANAGER

NOZZLE / E. L. DIENL................ o .... . ................. ° ........................................... ° ....... o ............................

15. RESPONSIBLE PROJECT ENGINEER (NAME / ORGANIZATION )D. J. _GNER / NOZZLE�FLEX BEARING PROJECT ENGINEERING

........... . ...................... ° .... ° ......................... . ......................... . ................. . ..........

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORr..,ANI_TION)

S. A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING

17. DESCRIPTION (ATTACH PFOR, FI_RES, PHOTOGRAPHS, ETC.)ALuminumoxiclecorrosionwas observecIbetweenthe primary and secondar"yo-rings, and outboard of the secondaryo-ring in the 360LOOIA (LN) aft exit cone field joint. No pitting was observed. Structural Al:_Lications inslccH:torsnoted that grease in both joints appeared lighter than required on STg7-2999,

Minor corrosion has I_ observed within past flight aft exit cone field joints, but has not been d_umented.

............................ • ......... ° ..................................................................... ° .... ° .....

18. JUSTIFICATION OF CLASSIFICATION (POSTFIRE ENGINEERING EVALUATION LIMITS)

This is not a flight issue. It is loeLieve<i that sea water entered past the primary o-ring at splashdown, resulting

in the minor corrosion observed. This is potentially a reuse issue if the grease application specification

(STW7-29_x_) is not followed........... • .... . ................. . .... ° ........................................................ • ............ ° ..........

19. CAUSE

Sea water entering past the primary o-ning at splashdown.

20. RECOMMENDED CORRECTIVE ACTION

1. Train arct certify SPC operators on grease aDpl(cation parSTW7-2999.

........... ° .... ° .... ° ............ o .....................................................................................

21. ANOMALY APPROVAL SIGNATURERPRB SECRETARY: DATE:

/S/T. L. JOHNSON 11/10/88

2. Do not write this minor corrosion up as a PR in the future.2Z. OB._RVATION/_NCW'4ALY APPROVAL SIGNATURES

PE:

................ ° .... ° ....... ° .... , .... = ................. . ....... = ......................................................

23. RESULTS OF RECO_ENDED CORRECTIVE ACTION1. Effective

2. Minor corrosion has been observed in this joint on the aft exitcone field joints of 360L002 and 360L003. Observed grease

al_oLications were nominal. This was not written uko as an observation.

26. REPORT RESULTS TO RPRB? YES X NO

................ . ................. ° ...........

25. RPRB CLOSURE SIGNATURE

(REQUIRED ONLY IF BLOCK 24 CHECKED "YES'*)RPRB SECRETARY: DATE:

26. OBSERVATION/ANOMALY CLOSURE SIGNATUREPM: DATE:

........... ° ............ ° .............. _ ....................................... • ............ ° ............................

27. ORIGINATION DATE I 28. REQUIRED STATUS DATE I 29. PR CLOSURE DATE J 30. PFAR CLOSURE DATE11109/88 J 11107188 J 01/1Z/B9 I

...................... °.°° .... o ......... . ......................... • ........................................... • ...........

REV. 3/28/89

ORIGINAL PAGE IS

OF POOR QUALITYTWR-17272B-2


1. PFAR NUMBER 3. INSPECTION LOCATION 4. REFERENCE SQUAWK NUMBER 5. REFERENCE PR _UMSER

360LOO1A-12 KSC X T.24/T-97 26-0046 , PV6-t11290...... . ............... . ............................. , ........................................

Z. aRM MOTOR HUMBER H-7 A-2 6. REFERENCE [FA _UMgER 7. _EFERENCE SPR NUMBER

360LOO1A ...... I _/A _/A

8. TITLECORROSION SETWEEN PRIMARY AND SECONDARY AFT EXIT CONE O-RINGS

...... ......o ......... . ........ • ...... . ...... • ............... . .................................................... . .....

9. CLASSIFICATION

OBSERVATION MINOR ANOMALY X MAJOR ANOMALY CRITICAL ANOMALY.... . .... ...

...o...... .... o•...o.........o.• .... ... ...... • ............... • ........ . ....................... .°.o. .....................

lo. PARTN_SER I 11. SERIALNUMBER I _Z. PARTDESCRIPTION1USZe39-01 I O00000S I FORWAROeXIT CONEASS_

... ..... . ..... ....... ...... . ....... °o. ............... • ............... • ........ . ...... . ...... . .........................

13. REPORTED BY (NAME /ORGANIZATION/ OOSERVATIOR DATE]K. B. NIELSEN / NTI QA / 10/0_88

....... • ...... °..° ..... . ...... . ...... . ............... . ........................ . .................................... •°.

16. RESPONSIBLE COMPONENT TEAN / PROGRAN MANAGER

HOZZLE / E. L. DIEHL....... . ...... °°. ...... ° .................................................................... ° ...... . ................. ,

15. RESPONSIBLE PROJECT ENGINEER iNANE / ORGANIZATION )

D. J. WAGNER / HOZZLE/FLEX BEARING PROJECT ENGINEERING....... °...o..... ...... °** .... _ .... .°•.°..... ............... .......... ...................... • ........................ ..

16. RESPONSIBLE DESIGN ENGINEER (NAHE / ORGANIZATION)S. A. NEYER / P_OZZLE/FLEX 8EARING DESIGN ENGINEERING

.......... . ...... . .... ... ...... . .......... . ........... . ............... . ............................. . ...................

17. DESCRIPTION (ATTACH PFOR, FIGURES, PHOTOGRAPHS, ETC.)

Aluminulloxidecorrosionwas observed between the primary and secondary o-rings, and outboard of the secorclary

o-ring, in the _L001A (LH) aft exit cone field joint. No pitting was _erv_. Structural apc_licatimsi_l:_tors _tN that grease in _th joints al=c_ear_ lighter than r_ir_ _ ST_7-29_.

Ni_r oorrosi_ has be_ ob_lerved _ithln l:_at flight aft exit c_ fleld joints, but has _t _n _t_.

.......... • ...... • ...... ° ...... °o° ......................................................................................

18, JUSTIFICATION OF CLASSIFICATION (POSTFIRE ENGINEERING EVALUATION LINITS)

This is not a flight issue. It is believed that sea water entered past the primary o-ring at splashdot_n, resultingin the minor corrosion observed. This is potentially a reuse issue if the grease application specificatio_(5TW7-2999) is not follc,_ed.

.......... • ...... ° ...... ° ........ . ...... ° ...... o ...... . ........ o ............................................. . ........ •.

19. CAUSE

Sea _eter entering past the primary o-ring at splashdown.

o°° ..... ° ...................................... . .................................................................... °.o

20. RECOIe4ENDED CORRECTIVE ACTION 21. ANOMALY APPROVAL SIGNATURE1. Train and certify SPC operators on greaae application per RPRS SECRETARY: DATE:

ST_7-2999. /S/T. L. JOHNSON 11/10/88

2. Do not wr_te th{s m_nor corrosion up as a PR in the future. - ............................................

22. OBSF,RVATION/kNOMALY APPROVAL SIGNATURES

re: _. _. (_..'_l_,""r,-,Q...'-- OATE:&J_..")/_',_

.......... ooo .................. o ............... o .................................................... o ............... o.o.

23. RESULTS OF RECOMNENDED CORRECTIVE ACTION 26. REPORT RESULTS TO RPRB? YES X NO1. Effective . - .....

2. Ninor corroSion has been observed in this joint on the aft exit 25. RPRB CLOSURE SIGNATURE

cone field joints of 360L002 and 360L003. Observed grease (REOUIRED ONLY IF BLOCK 26 CHECKED "YES")applications were nominal. This _aa not written up as an observation. RPRB SECRETARY: DATE:

26. OBSERVATION/ANOMALY CLOSURE S]GNATUREPI,I: OATE:

.................................... o .......... o .................................................... . ...................

27. ORIGINATION DATE I 2e° REOUIRED STATUS DATE I 29. PR CLOSURE DATE I 30. PFAR CLOSURE DATE

11/09/88 I 11/07/88 I 11/12/89 I

REV. 3/28189

TWR-17272B-3

POSTF[RE ANCNALY RECORD (PFAR)

............ o ................ " ................... ° .......................................... 4 ............. 4 ..............

I. PFAR NUMBER 3. INSPECTION LOCATION 4. REFERENCE SQUAWK NUMBER 5. REFERENCE PR NUMBER

360LOO1A-63 KSC T-26/T-97 N/A N/A........... . ............. _°.. ........ . ........................................................

2. SRM MOTOR NUMBER N-7 X k-2 6. REFERENCE IFA _UNBER 7. REFERENCE SPR _UNBER

]60LOO1A ...... M/A N/A........... 4 .... " ...... "°_ ........... °°4°°" ..... ° ................... * .................................... 4.--" ...........

8. TITLERTV/EA913 ADHESIVE MIXING IN JOINT 2

........... °.... ...... .°.4°°. .......... .°-..°. ................ 4 ........ . ............. ° ................... ..°4 ...........

9. CLASSIFICATIONOBSERVATION MINOR ANOMALY X MAJOR ANO/4ALY CRITICAL ANOMALY

.. ..........

.......... o .... .°.....°.o... ........ .°l...o...°. ............. _ ........ 4 ............. . ..... . ............. 4°°. ...........

N/A C_L/NOSE INLET JOINT (JOINT 2)..... ........................ ......... ..°........ .... °°°.°....° ..... ..°. ............. ° ..... ..................... ........

13. REPORTED BY (NAME /ORGANIZATION /OBSERVATION DATE_R. J. GEORGE / NOZZLE/FLEX BEARING DESIGN ENGINEERING / 10110/88

........ . .......... . ........ ..... ...... ......° .......... ......4...°.. ...................... 4.... ......... . ..... . ........

16. RESPONSIBLE COMPONENT TEN4 / PROGRAM MANAGERNOZZLE / E. L. DIEHL

15. RESPONSIBLE PROJECT ENGINEER (NAME / ORGANIT.ATION )O. J. WAGNER / NOZZLE/FLEX BEARING PROJECT ENGINEERING

..... ...4... ...... .. ................... .°..4.4.....4 .......... . ..... .°.. ............. .... ......... .. ..... 4...... ........

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORCu_NIT,ATION)

S, A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING........ = ............. ---- ............. ------- .......... - ..... o--- ................... --o---- ............. ° ..............

17. DESCRIPTION (ATTACH PFOR, FII3JRES, PHOTOGRAPHS, ETC.)See continuationsheet.

..... 4°°4.°" .... "°°Q°''''" ....... " ..... 0°''''" ............. 4 ..... ..°. ...... .. ........ °.°° ................ ..°. ...........

18, JUSTIFICATION OF CLASSIFICATION (POSTFIRE ENGINEERING EVALUATION LIMITS)

There is no aOheeion bet_.en the RTV and EA913 adhesive, reducing the capai=i[ity of the RTV to act as a thermalbarrier. The eroded EA913 NA edhesive has not previously been observed (significant departure from the historical

data base)........... °. ....... .°..°.. .... .°°.°...°.... ........ . ....... . ................ . ...................... .°°.... ..............

19. CAUSECurrent bottling and joint assembly procedures.

..... _ ....... ..°.°.°....°.°° ..... . ..... .°.....o.°.°...°......°. ............. ° ..... .°.. ............. °°..°.. ....... .°...°4

20. RECOI4_IENDED CORRECTIVE ACTION 21. ANOMALY APPROVAL SIGNATURE

At the time this us presented to the RPRB, the PV-1 nozzle Joint #2 RPRB SECRETARY: DATE:

had just been exmined. It _as recommended that we further evaluate , r,_, ¢_ 3"/c/c;"the new assemOLy procedure change on QN-8. It would then be

determined if the neu asselmly process would be incorporated into ........ 1 ..... y .........................

flight nozzleS. 22. OB.T_E.RVA'I_,zO.N,gANONALYAPPROVAL _IGNATURES

PE_/_/,_ _¢_ _,_ DATE: _-,/--_ "_

23. RESULTS OF RECONNENDED CORRECTIVE ACTION 26. REPORT RESULTS TO RPRB? YES X NO

See continuation sheet. - ......_......... .... ...°.......°. .......... °......


(RE_JIRED ONLY IF BLOCK 26 CHECKED "YES")RBRB SECRETARY: _ DATE:

........ -;. ..... _ ............... _- ...........Z6. OBSERVATIOM/.ENONAI.3 CLOSURE SIGNATURE

°...°.°°...°°..°_°.. ....... .....°. ....... . ..... ....°...°..°. .......... 4.°..°.°..°°. ........... .°.°.. .......... .-°. ......

..°_ ...... .°... ...... . ...... . ..... . ........... °°.°..... ................ °°..°°.... ................ 4... ...... _ ...... ° .......

REV. $/Z8/89

TWR-17272B-4

CONTINUATION SHEET FOR PFAR NUNEER 360LOO1A-63

17. DESCRIPTION (continuation)

The EA913 adhesive used to bond the cowl insulation to the cowl housing extrucied into the radial RTV bondLine

between the cowl tin9 an<:lnose cap 360-d_. oircumferentiatly. The a_=hesive that extr_ed into the raaial b_ndtine

was typically saOiched between t_o layers of RTV. Soot was observe_ over a majority of the joint circumference

and up to the primary o-ri_. A distinct blowpath was observed at 216 dq. Heat affected GCP ar_:lSCP _ere o_erv_,_:l

at the bLo_th location. The charred _teria( was aporoximately 0.01 in. deep. The EA913 NA _esive adjacent to

the cowl housing was erred a_roximately 0.1 in. axially x 0.7 in. wide circ. at the blow_th location. There was

no bloody, erosion or heat effect to the printmry o-ring. Electrical c_tivity measurements on the coul ar_ nose

inlet housings sho_ed no heat damage.

The mixing of the adhesive in the co.t/nose cap bondline and the presence of soot was documented on ETM-1A, DM-6,DM-9, OM-6 and ON-7. It is believed that the mixing has occurrea on all previous nozzles due _o assemDLy

proceoures. Charred GCP and SCP insulators have been observed at btowpath Locations in the ON-6, and PV-1 nozzleinternal Joint #2. Erocled EA913 HA aa_osive has not previously been observed.

23. RESULTS OF RECOMMENDED CORRECTIVE ACTZON (continuation)

1. The new assembly procedure _ms tested on PV-1 and QM-8 Joint #2's. Post-test inspections of these jointspromoted the decision to further evaluate and refine the ne_ asselmty procedure on TEN motors.

As of 360L002, there has never been Drimmry o-ring blo_nf, erosion or heat effect, or metal heat damage with Joint

#2. In the 360L002 RPRB, it was decided to close out all "minor a_lies" similar to this on all previous static

tests and flights. ALL future occurrences, fncl_i_ 300L002, _ld be ct_nsidered "observations,%

TWR-17198 Vo(. V has been uodated to include accept/reject criteria for RTV in nozzle joints. The condition notedin this "minor anomaly" is now an "acceptable" condition per this criteria.

TWR-17272B-5

POSTFIRE ANC_ALY RECORD (PFAR)

.................... • oo ...... • .......... • .................... .oo ............... ° ............... o ...... • ................. _.

1. PFAR NUMBER 3. INSPECT]ON LOCATION 4. REFERENCE SQUAWK NUMBER _. REFERENCE PR NUMBER

360LOOIB-IO KSC X T-26/T 97 26-0043 PV6-111312.o .............................. , .............................................................

2. SRM MOTOR NUMBER H-7 A-2 6. REFERENCE IFA NUMBER 7. REFERENCE SPR NUMBER

36GLOOIS ...... I N/A I NIA................... o.• ...... , ...........................................................................................

8. TITLECORROSION BETWEEN PRIMARY AND SECONOARY AFT EXIT CONE O-RINGS

................... ..• ........ . ...... ,.• ............. , ...................................... . ...........................

9. CLASSIFICATION

OBSERVATION MINOR ANOMALY X MAJOR ANOMALY CRITICAL ANOMALY.. .... .o. ...

.oo. ...... •.... .... ...o..oo., .... ...., .... .°..• ...... . ...... .....o..o, ........ o ............................... ° ........ o

10, PART NUMBER J 11. SERIAL NUMBER I 12. PART DESCRIPTION

1U76039"03 I 0000001 I AFT EXIT CONE ASSY................... ..o.. .... . .......... o ............................. . ...................... .o° .........................

13. REPORTED BY (NAME / ORGANIZATION / OBSERVATION DATE)K. B. NIELSEN / MT! QA I 10107188

.......... • .... .....o. ............... . ............................... . ........................................ . .........

14. RESPONSIBLE COMPONENT TEAM / PROGRAM MANAGERNOZZLE / E. L. DIEHL

.° ................. •.. ...... . ........ ° .... . .......... ° ........ o .................... o ........ ..° ...... , ........ . ........ °

15. RESPONSIBLE PROJECT ENGINEER (NAME / ORGANIZATION )

O. J. WAGNER / NOZZLE/FLEX BEARING PROJECT ENGINEERING................... • ........ . ........................ ° ...... •=° .... o ........ • ......... o ..... , ........................ ...

16. RESPONS!BLE DESIGN ENGINEER (NAME / ORGANIZATION)S, A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING

........ • .......... ° ..... -o-o ........ • ...... . ........ ° ...... .., .... , ........ ° ...... . ........ . ........ • ............... °°°


ALuminum oxide corrosionwaso_ervedbetweenthe primary ana secor_ry o-rings, and outboard of the secondaryo-ring, in the 360LOO1B (RN) aft exit cone fie{d joint. No pitting was observed. Structural Applicationsinspectors note¢_ that grease in both joints tpeared lighter than required on STWT-2999.

Minor cor-osion has been observed within past flight aft exit cone field joints, but has not been documented°

........ • ........ °=. ............... • .......... • ............. • ............... . ...... • ................. • ............... ..•

18. JUST%FICAT!ON OF CLASSIFICATION (POSTFIRE ENGINEERING E'VALUATION LIMITS)

This is not a flight issue. [t is believed that sea water entered past the primary o-ring at splashdown, resultingin the minor corrosion observed. This is potentially a reuse issue if the grease application specification(STWT-2999) ia not fo(lo_e¢l•

................ • .• ............................................... . ........ o ...... o ........ • ........................ °..

19. CAUSE

Sea water entering paat the primary o-ring at splashdown.

.................................... ° ...... o ............................... • .......................................... o

20• RECOMMENDED CORRECTIVE ACTION 21. ANOMALY APPROVAL SIGNATURE

I. Train and certify SPC operators on grease application per RPRB SECRETARY: DATE:

STW7-2999. /S/T. L. JOHNSON 11/10/88

2. Do not write this minor corrosion up al s PR in the future. - ............................................

2p21_: O_R_AT L%__PPROVAAL TS ! _N_

oo ......... . ..... •.. .... •.. ................. • ........ . ...... • ...... . ........ • ...........................................

23. RESULTS OF RECOMMENDED CORRECTIVE ACTION 24. REPORT RESULTS TO RPRB? YES X NOI. Effective

2. Minor corrosion hal b_ observed in this joint and on the aft

exit cone field joints of _._A_LO02 and 360L003. Observed greaseapplications were nominal. This was not written up as an ot_ervetion,

..... o

o• ...... o ....................................


(REQUIREO ONLY IF BLOCK 24 CHECKED "YES")RPRB SECRETARY: DATE:

26. OBSERVATION/ANOMALY CLOSURE SIGNATURE

PM: DATE:

................. oo• ........................................ • .......................................................... o

_7" OR I i_"I NAT !ON_)ATEII/09/88 I 28° REQUIRED STATUS DATE I _9" PR cLOSUREDATE!I/07188 01 / 12/89 J _0. PFAR _LOSURE DATE

.................. • .......................................................... • ...... • ................................. •ooo

REV. 3/28/89

TWR-17272B-6


.. ......... . ......... . ......... o ...... .o.oo.. ............. .... ............................................................

I. PFAR NUMBER 3. INSPECTION LOCATION 4. REFERENCE SQUAWK NUMBER 5. REFERENCE PR NUMBER

360LOOIB-38 KSC X T-E4/T-97 26-0048 _ PV6-11129Z.......... o ..................................................................................

2. SRM MOTOR NUMBER N-7 A-2 6. REFERENCE IFA NUMBER 7. REFERENCE SPR NUMBER

360LOO1B ...... I N/A N/A.............................. • ...... . ...................... • ...........................................................

8. TITLEDING ON F_JD EXIT CONE AFT FLANGE AT 97.6 DEG

.......... • ......... . ................... . ........................................................ . ......................

9, CLASSIFICATIONOBSERVATION MINOR ANOMALY X MAJOR ANO?4ALY CRITICAL ANOMALY

.................... . ....................................... . ...... . ......... . ..........................................

_o. PARTN_BER 11-OOO06 _2. PARTDESCRIPTZON1U52839-01 FORWARD EXIT CONE ASSY.......... ...o. .......... o .... . ......... . ...... .. .................. = ....................................................

13. REPORTED BY (NAME / ORGANIZATION / OBSERVATION DATE)

T. B. GREGORIO / MT% QA / 10/07/88.......... o ................... o .................................................................. o ......................

14. RESPONSIBLE COt41>(]NENT TEAM / PROGRAM MANAGER

NOZZLE / E. L. DIEHL.......... o ................... . ......... ° .......................... o ................... . ......... ° ......................

15. RESPONSIBLE PROJECT ENGINEER (NAME / ORGANIZATION )O. J. WAGNER / I_ZZLE/FLBX BEARING PROJECT ENGINEERING

.......... • ...... .=.. ...... .... ........................................................ o ...... .=.° ...... . ..............

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORGANIZATION)S. A. MEYER / NOZZLE/FLEX SEARING DESIGN ENGINEERING

.......... o ......... . ....... ... ......... .°.o. ............ ..o. ...... o ......... ° ................ ° ............... . .........


During the 360LOO1B (RH) eft exit corm dmte, I guide p(n dfnged the edge of e hole on the forward exit cone aftflange. The dine was ap_roximately 0.02 incn_ deep.

Damage his occurred during KSC aft exit cone demates on previous flight nozzles. SRM-15A and 23A forward exit coneaft flange seating surfaces were scratched by the joint alignment pin. The 360LOOIA (LH) forward exit corle aftflange was also scratched 0.002 inches deep during demate bye guide pin.

.......... . ...... o.°....°.°.o°..° .... °.°°.°°°°.. ............ ° ..... o ..... ° .... ° ...... °o°o.°°°°..°.. ...... . ..............


The current KSC aft exit col_ decmting procedure/tooling could cause dammge preventing reuse of hardware.

Requires corrective ection but has no impact on motor performance or program schedule....... .... .... ... ........ o.... .... o.. ......... . ................... ° ..... ..... .......................... . ..............

19. CAUSE

Current KSC aft exit cone dmting procedures and tooting.

20. RECOI4NENOED CORRECTIVE ACTION

See continuation sheet.

...... °°.° ................ °.°o ...... _..° ...... ° ....................................... ° ...... o .........................

21. ANOMALY APPROVAL SIGNATURE

RPRB SECRETARY : DATE :

/S/T. L. JOHNSON 11/10/8a

22. OBSF,_VATION/ANOMALY APPROVAL SIGNATURE3

PE: _,_. _J'_L,¢_ DATE: b/_.'_/_

............. . ............. .... ...... oo.. ...... . ............ . ...........................................................

23. RESULTS OF RECOMMENDED CORRECTIVE ACTION 2(*. REPORT RESULTS TO RPRB? YES X NO

See continuatiotl stleet. - ........ ......... ° ...... .... .....................

25. RPR8 CLOSURE SIGNATURE(REQUIRED ONLY IF BLOCK 2{* CHECKED "YES")RPRB SECRETARY : DATE :

26. OBSERVATION/ANOMALY CLOSURE SIGNATUREPM: DATE :

....... o.oo ...... . .......................................... o ...... o .......................... o .........................

27. ORIGINATION DATE I 28. REQUIRED STATUS DATE I 29. PR CLOSURE DATE I 30. PFAR CLOSURE DATElI/09/88 I 11/07/88 I _1122188 I

........ o... ...... ° .......................................................................................................

REV. 3128189

TWR-17272

B-7

CONTINUATION SHEET FOR PFAR NUMBER 360LOOIB-_8

20. RECONMENOED CORRECTIVE ACTION (continuation)

I. ELiminate guide pins. These just add more Oossibltit_es of scrstchlng the forward exit cone aft flange after

the joint serrates an recoils. ELiminating the guide pins leaves only the joint alignment pin to worry amour.

2. Wear term: Use anti-r_,4_oil toots to eliminate the recoil of the aft exit cone after joint se!_wlrstion. Thiswould not allow the atignfaent pin to contact the forward exit cone aft flange.

ASAP: Design and incorloorste a rail system disasseelioly tool similar to that used at MTI Wasatch T-24 arx:l T-97.

This tool su;:_oorts the aft exit cone assenibly so that there is no load on the joint alignn_t Din during seoaration.The tool also do. not recoil, eliminating any possibility of damage to the forward exit cone aft flange t_y the

aligrnent pin.

23. RESULTS OF RECOMMENDED CORRECTIVE ACTION (continuation)

I. Effective: RSRN Flight Z (360L002)

2. Anti-recoiL tooLs effective:

RaiL system tool effective:

KSC has incorporated a toed cell into the existing aft exit cone stub removaL tool. This Load ceil allows operatorsto monitor the load that is being suDDorted by the fork lift. If used correctly, the load on the joint aLigr_t

pin can be minimizad during seoaration. This would eliminate relative vertical dispkacen_-nts between the aft exitcone anct forward exit cone after joint set:station. The load cell was used on the aft exit cone Oemates of RSRM

Flights 2 and 3 (360L002 and 360L003). Only minor raisad metal was oioserved on the forward exit cone 91.8 degreealignment pin holes on the four joints where the load cell was usad. This does not violate refurbishment

requirements.

TWR-17272

B-8

POSTFIRE ANC_JALY RECORD (PFAR)

... .................................. .o ......... . ....... ..... ................. . .... .... ....... . .... . .................. ....

1. PFAR NUMBER I 3. INSPECTION LOCATION _ 4. REFERENCE SQUAWK NUMBER 5. REFERENCE PR _UMBER

360LOOIB-42 1 KS( X T-24/T-97 J 26-0044 _ PV6-111324

. ....... o... ....... . ....... o ....................... . .............. ° .... . .....................

2. SRM MOTOR NUMBER H-7 A-2 _ 6. REFERENCE IFA NUMBER 7. REFERENCE SPR _UMBER360LOD1B ...... _ N/A N/A

8. TITLECORROSION BETWEEN PRIMARY AND SECONOARY FORWARD EXIT CONE O-RINGS

.. .... o..oo.. ............ . ....... o... ........ ... ....... . .......... o .................. . ............ . .....................

9. CLASSIFICATION

OBSERVATION MINOR ANOMALY X MAJOR ANOMALY CRZTICAL ANOMALY

............ . ...... °. ....... . ....... o .......... . ............... o .................................. . .....................

lo PARTNUMBER I_1 SERIALUBER 1_2 PARTDESCRIPTZOR1U52839"01 I 0000006 J FORWARD EXIT CONE ASSY

.o ....... °o.. ....... . ......... . ..... ..°. ....... . ....... .o.. .............................................................

13. REPORTED BY (NAME / ORGANIZATION/OBSERVATION DATE)K. B. NIELSEN / MTI QA / 10/0_88

.. ....... . .......... . ....... . .................. . ....... . ............. . ..................................................

14. RESPONSIBLE COMPONENT TEAM / PROGRAM MANAGERNOZZLE / E. L. OlENL

.°o°. .... °o.. ....... ° ....... . ....... .. ......... ..o.o....... ....... o..° ....... ° ....... ° ............ . .....................


O. J. WAGNER / NOZZLE/FLEX BEARING PROJECT ENGINEERING.... .... .... . ..... ..° ....... . ....... .... ....... ° .......... . ....... ........o.=. ....... ......o.. .... . ............. . .......

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORGANIZATION)

S. A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING......... .°.. ............... . ....... . ....................... . ........... . .... . ....... . ............ . ............. o .......

17. DESCRIPTION (ATTACH PFOR, FIGURES, PHOTOGRAPHS, ETC.)Al_in_ oxide corrosi_ was observed loetw_ the primary and s_ondary o-rings, and outboard of the secoc_ry

o-ring, in the 360LOOIB (RH) aft exit cone field joint. No pitting was o_ervea. StructurmL Applicationsinspectors noted that grease in both joints appeared lighter than required on STW7-29Q9.

Minor corrosion hal been ob6erved within past flight aft exit corNefield joints, but has not been documented.

........................... ° ....... o°o° ............... . ............. ° ....... ° ....... ° ............ ° ............. ° .......


This is not a flight issue. It is believed that see water entered past the primary o-ring at splashdown, resultingin the corrosion observed. This is potentially m rue issue if the grease al_olication specification (STW7-2999) isnot fol Lowed.

ooo. ....... o ....... o ...................................................................................................

19. CAUSE

Sel water angering past the primary o-ring at splashdown.

......... oo.o ....... o ..................................... o .................. o ............................... .o.o .......

20. RECOMMENDED CORRECTIVE ACTION 21. ANOMALY APPROVAL SIGNATURE

1. Train end certify SPC ol_ratorl on greaee application par RPRB SECRETARY: DATE:STW7-2999. /S/T. L. JONNSON 11/10/88

2. 0o not write this minor corrosion up mm a PR in the future. -............................................

22. OB_R_(ATION/_NOMALY APPROVAL SIGNATURES

=.°.. .... o... ....... o ....... o ....... o... ....... .=....... ..................................... . ..........................

;[3. RESULTS OF RECOMNIENOED CORRECTIVE ACTION 24. REPORT RESULTS TO RPRB? YES X NO1. Effective ......

2. Minor corrosion has been observed in this joint on the aft exitcone field joints of 360L002 and 360L003. Observed grease

applicationl were rmminet. This wee not written up el an o_=ervetion.

.......... o ............ o .................. oo.


(REQUIRED ONLY IF BLOCK 24 CHECKED "YES")RPRB SECRETARY: DATE:

26. OBSERVATION�ANOMALY CLOSURE SIGNATUREPM: DATE:

.... o ....... o ....................... oo.o ....... o .......... o .......................... . ....... o ..........................

27. ORIG|NATION DATE _ 28. REQUIRED STATUS DATE I 29. PR CLOSURE DATE I 30. PFAR CLOSURE DATE11/09/88 'l 11107/88 ) 01/12/S9 I

............. . ....... o. ...... °... .... ° ................................................ , ...................................

REV. 3/28/89

TWR-17272B-9


............................................................................ o .............................................

I. PFAR _UMBER I 3. INSPECTION LOCATION 4. REFERENCE SOUAWK NUMBER 5. REFERENCE PR NUMBER

360LOOIB-_4 1 KSC T-24/T.97 _/A N/A

............................................ .=oo. ........... • .................................

2. SRM MOTOR _L94BER H-7 X A-2 6. REFERENCE ]FA _UMBER 7. REFERENCE $PR NUMBER360LOOIB ...... _/A N/A

.................. • .o...°oo.. ....... o ...................................................... . ............................

8. TITLERTV 8ACKFILL IN JOINT _ INHIBITED SY EXCESSIVE GREASE

.................. o ........... ..°.• ..... o ...............................................................................

9. CLASSIFICATION

OBSERVATION MINOR ANI_ALY X MAJOR ANOMALY CRITICAL ANOHALY• ..... .°° ...

.................. • o... ..... . ........................................ o ................. • ................................

10. PART NUMBER 11. SERIAL NUMBER 12. PART DESCRIPTIONN/A I NIA 1 NIA

.................. o...• .............................................................. o ..................................

13, REPORTED BY (NAME / ORGANIZATION / OBSERVATION DATE)

R. J. GEORGE / NOZZLE/FLEX BEARING DESIGN ENGINEERING / 11/03/88.................. o.°.ooo.o.o ..... .=o.. .................................... • ............... .=. ....

14. RESPONSIBLE COMPONENT TEAM / PROGRAM MANAGER ......................

NOZZLE / E. L. DIEHL...................... • .°.o.° ..... = .................................................... •°o.. ...........................

15. RESPONSIBLE PROJECT ENGINEER (NAME / ORGANIZATION )O. J. WAGNER / NOZZLE/FLEX BEARING PROJECT ENGINEERING

.................................. • .............................. • ........................... =.. ........................

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORGANIZATION)

S. A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGrNEERING.................. • ..o° ........................................... o ...... ... ...........................................

17. DESCRIPTION (ATTACH PFOR, FIGURES, PHOTOGRAPHS, ETC.)See continuation sheet.

.................. °.°.°°°=o°o .................................... o ....... • ......... o°° ........ ..........................

18, JUSTIFICATION OF CLASSIFICATION (POSTFIRE ENGINEERING EVALUATION LIMITS)See continuation sheet.

o,°o ........ ° ..... °=ooo ..... *°.-*°q.°°.-°-. ................... o°°*°oo° ............. °o°°.....°° ..........................

19. CAUSE

Excess grease applied during joint assembly resulted because no application standards were in place°

.................. .°°oo°°o°.o.ooo ......... . .............................................. o ..............................

20. RECOMMENDED CORRECTIVE ACTTON Z1. ANOMALY APPROVAL SIGNATURE1. Set up TRACS class to train and qualify technicians, QA and AF on

grease appticatioq_.

2. Continue inspecting future post-test and flight nozzle joints for

excess grease problems.

3. Release and Incorporate Rev. C to STMT-Z999.

RPR8 SECRETARY : DATE :

............ ..o ..... .°° ......................

22. OBSERVATION/ANOMALY APPROVAL SIGNATURESPE: DATE:

PM; DATE :

.................. ° ......... °.°o°o•°°.o .................................... o ............. o°o ..... ° ......................

23. RESULTS OF RECOMMENDED CORRECTIVE ACTION Z4. REPORT RESULTS TO RPRB? YES X NOSee continuation sheet. - .....

........ °...°...o ............................

ZS. RPR8 CLOSURE SIGNATURE

(REOUIRED ONLY IF BLOCK 24 CHECKED "YES")RPRB SECRETARY: DATE :

.=.°• ....... •=oo° ............................

26. 08SERVATION/ANO_ALY CLOSURE SIGNATUREPM: DATE:

.......................... °o°oo .......................................................... • ..............................

27. ORIGINATIONDATE I 2e. REOUIREDSTATUSDATE I zg. Pe CLOSUREDATe I SO. PFARCLOSUREDATE11/3o/_ I N/A I N/A I

REV. 3128189

TWR-17272B-IO

CONTINUAT_OH SHEET FOR PFAR NUMBER 360LOOIB'44

17. DESCRIPTION (continuation)

The S60LOOIB nozzle forward exit cone/throat joint (Joint _z.) showed excessive grease. Grease was observeQ on the

radial O.O. and on the axial portions of the joint interface. It is believed that the oresence of the grease at 185

degrees inhibited the backfill of RTV, A bLowpath was Located at the 185 Oegree location. There was no OLowoy,erosion or heat effect to primary o'ring.

Joint #4 on the DM-9, QM-6, QM-7 and PV-1 tests showed excessive grease on phenolic interfaces. [t was believedthat the excess grease inhibite_l the RTV backfill in these joints (classified as Minor Anomalies).

Previous flight and static test forward exit cone/throat joints have shown grease on the _enolic joint interfaces.

[n order to hold the primary o-ring on the o-ring groove during assemoty, a thicker layer of grease is applies tothe o-ring and groove.

STgT-Z999 was released to control amounts of grease al_liecl to the nozzle internal joint o,rings and o-ring grooves.This specification was dOSed to engineering assembly drawings and will be effective RSRM flight 6 (360T004).

18. JUSTIFICATION OF CLA$SIFICATION (continuation)

The excess grease in joint #4 recluired corrective action but has no impact on motor performance or program schedule.

The excess grease inhibited the RTV backfill at the 185 degree location.

Grease on the joint phenolic interfaces could prevent the adhesion of the RTV to the phenolic. This would re_uce

the capability of the RTV to act as a thermal barrier.

25. RESULTS OF RECOMMENDED CORRECTIVE ACTION (continuation)

STW7-2999 was incorporetecl into nozzle joint assembly planning; effective QM-8 and RSRM flight 4 (360T004).

1. Effective

2, QM-8 showed excess grease inhibiting RTV backfill in joint #4. The grease ac_Lication specification did not

reduce the excess grease in this joint. Joint #4 will be monitored by structural applications until proper greaseapplications are being followed. Also, TRACS classes are being set up to train/qualify nozzle joint assermoLypersonnel (effective ).

S60LOO1A showed no excess grease in nozzle joints.

360LOOZA showed no excess grease in nozzle joints.

360LOOEB showed no excess grease in nozzle joints.

360LOO3A showed no excess grease in nozzle joints.

360LOO3B showed no excess grease in nozzle joints.

TWR-17272B-II


1, PFAR NUMBER 3. INSPECTION LOCATION _. REFERENCE SQUAWI_ _U_RER 5o REFERENCE PR _UMBER

36plOD. T2 ,T. I,!: N,A............................. . ........ o ........ ....oo..

2. SRM MOTOR NUMBER H-7 X A-2 _ 6. REFERENCE IFA NUMBER 7. REFERENCE SPR NUNBER360LDO1B ...... _ N/A _/A

..o ........ .o.o..oo ....... ... ........ . ..................... . ........ . ..................... o.... .........................

8. TITLERTVIEA913 ADHESIVE MIXING IM JOINT 2

......... o ..... ... .... ....... ........ . ........ . ............ . ............ . ................. ..... ........ . ................

9. CLASSIFICAT[ON

OBSERVATION MINOR ANOMALY X HAJOR ANOMALY CRITICAL ANOMALY...... ooo o..

..o ........ .o.o.oo.o .... ..... ........ o ..................... . ........ ..... ..................... . .........................

10. PART NUMBER I 11. SERIAL NUMBER I 12. PART DESCRIPTION

N/A I ./A I CO_L/_OSEINLETJOINT(JOiHT2)o.......o. .... ..... ..... ..................... ......... . .... o ........ o.o.. .... .. ........... ..o.. ............... .o=...o...

13. REPORTED BY (MANE / ORGANIZATION / OBSERVATION DATE)S. A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING / 11/04/89

............... . ...... ....... ........ o... ..... ... .......... . .............................. ..... ........ . ................

14. RESPONSIBLE C_41_3NENT TEAM / PROGRAM MANAGER

NOZZLE / E. L. DIEHL... ........ ...o. ........ ...o. ........ ...o. ................. .. ............................. ..... ..................... o.


O. J. WAGNER / NOZZLE/FLEX BEARING PROJECT ENGINEERINGo.= ........ ...o.........o.=.o ........ o .... . ............. .oo.oo.. ........ . ..................... .. ....... .... ..... ........

16. RESPONSIBLE DESIGN ENGINEER (NAME / ORGANIZAT%ON)

S. A. MEYER / NOZZLE/FLEX BEARING DESIGN ENGINEERING... ............ ..... .... ...._ ........ . ....... . ............. . ............ . ............... ......_ ........ _ ........ ........

17. DESCRIPTION (ATTACH PFOR, FIGURES, PHOTOGRAPHS, ETC.)The EA913 aOhesive used to bond the cowl insulation to the cowl housing extruded fOrD the radial RTV bondtinebetween the couL ring and nose cap 360-deg. circumferentiakly. The aOhesive that extrucJed into the r_liaL bondline

was typically sanOwiched between two Layers of RTV. Soot was observed over a ,mjority of the joint circumference

ana up to the axial bolt holes. There was no bloutw, erosio_ or heat effect to the primary o-ring.

The mixing of the ac_esive in the coul/nose cap bondline and the presence of soot was documented on ETM-1A, DR-S,ON-9, ON-6 and _N-7. It is believed that the mixing has occurred on all previous nozzles due to assemblyprocedures.

o.o ...... .......o ........... o ........ . ............ . ........ ..... ........ o ........ . ............ . ................... *.o...

18. JUSTIFICATION OF CLASSIFICATION (POSTFIRE ENGINEERING EVALUATION LIMITS)There is no a_i_esion between the RTV and EA913 aahesive, reclucing the capabitity of the RTV to act as a thermalbarrier.

........... . ...... oo........o ........ . ........... ** ........ . ............ * ........ . ............. . ....... . ................

19. CAUSE

Current bonding and joint assembly procedures.

..o ............ _o..o ...... ... o.... o....o..**.* o..*..o._ ........ .o. ...... . ........ _o... ........ _. .......... o .............

20. RECOMMENDED CORRECTIVE ACTION 21. ANOMALY APPROVAL SIGNATURE

1. At the time this was presented to RPRB, the PV-1 nozzle Joint #2 RPRB SECRETARY: DATE:

hed j=t been exmined. [ t was recoemended that . further eva,uatenew process _L _'-_-] ["-_./.ithe new asse_loty procedure change on ON*8. It would then be C_.-I " _ _ I/..x.zdetermined if the assecmty would be incorporated into ............. ; ............ _ .... ;-;---flight nozzles. 22. OBsJ_,_VATIOJr/ANOMALY APPROVA¢.=,IGNATURES

o...... ........ _....o......o..... ...... oo.o ....... ..... ........ o .... . ............ ..

23. RESULTS OF RE_NOED CORRECTIVE ACTION 24. REPORT RESULTS TO RPRB? YES X NOSee continuation sheet. - .....

........... ........ . ....... o .... ..... ........

_.._. RPRB CLOSURE SIGNATURE

(REQUIRED ONLY IF BLOCK .24 CHECKED "YES")R_P._I8SECRETARY: , DATE:

..26. ORSE'RVATIO_ANOMAL_ CLOSURE SIGNATURE

.......................... ...... ..o.o...o_._.o.. ...... o .... ....= ................ o..ooo ..... . ............... .o.._ ........

27. OR IGIRAT_O_ DATE11/30/88 I 2'" REQU|RED STATUB DATE I 29" PR CLOsuRE DATENJA N/A I 30 ° _)FAR cLO_JRE DATE

....... o ........ ..o..o.o. .... ..o.. ..... .o... .................... o ..................... o ........ ..... ..... ooo.o.o. ........ o

REV. 3128189

TWR-17272B-12

CONTINUATION SHEET FOR PFAR NUMBER 360LOO1R-45

2.3. RESULTS OF RECOMMENDED CORRECTIVE ACTION (continuation)

1. The new assembly procedure was tested on PV-I and QN-8 Joint #2's. Post-test inspections of these joints

prompted the decision to further eveluate and reffne the new assemOLy pnoceaure on TEM motors.

As of 360L002, there has never been primary o-ring btowby, erosion or hest effect, or metal heat damaqe within Joint#2. In the 360L002 RPR8, it was decided to close out all "minor anomalies- similar to this on all previous staictests and flights. ALL future occurrences, incluaing 360L002, would be considered ,,ol_ervstions,,.

TMR-17198 VoL. V has been updated to include accept/reject criteria for RTV in nozzle joints. The condition notedin this ,,minor an(_aLy" is no_ an '*acceptabLe '_ condition per this criteria.

TWR-17272B-13

Recipient

K. Baker

D. D. BrightF W. Call

M Clark

J DonatE L. DiehlT FrestonJ E. Fonnesbeck

R J. GeorgeS Graves

D. Harris

J. D. Leavitt

R. Loevy

S. A. MeyerG. Nielson

D. NisongerC. Olsen

S. Olsen

B. E. PhippsR. B. Roth

G. Snlder

R. K. Wilks

Print Crib

Vault

DISTRIBUTION

No. of Copies

1

1

15

1

1

1

1

1

1

1

1

1

1

1

1

1

1

11

1

1

1

5

5

Mail Stop

287L22

E05

427

L62A

El4

242

L62A

L62A

L52

L34

El4

L62A

L62A

811

M31

411EL35

L22

LlO

411

L23

K23B

K23E

TWR-17272

46

Flight Motor Set 360L001 (STS-26R) Final Report

Documents

Transcript of Flight Motor Set 360L001 (STS-26R) Final Report