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BLOCK II SOLID ROCKET MOTOR (SRM) 

CONCEPTUAL DESIGN STUDY 

CONTRACT NAS 8137295 

FINAL REPORT - VOLUME I 

Prepared for: 

GEORGE C. MARSHALL SPACE FLIGHT CENTER 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

Marshall Space Flight Center, Alabama 3581 2 

Attention: AS24D 

(kASr-CR-17sc50) 

(SI) CClCEPTCAL TESIGk’ STUGP, VCLUUE 1 

Pinal ReFort (Atlantic Research Corp.) 

bLOCR 2 SCL30 LCCEET BOTOB 

151 p CSCL 21B 

G3/2 0 

Submitted by: 

ATLANTIC RESEARCH CORPORATION 

7511 WELLINGTON ROAD 

GAINESVILLE, VIRGINIA . 22065-1699 

TR-PL-12126 31 December 1986 

ne9- 13489 

Unclas 

0063962 
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BLOCK II SOLID ROCKET MOTOR (SRM) 

CONCEPTUAL DESIGN STUDY 

CONTRACT NAS 8-37295 

FINAL REPORT - VOLUME I 

Prepared for: 

George C. Marshall Space Flight Center 

National Aeronautics and Space Administration 

Marshall Space Flight Center, Alabama 3581 2 

Attention: ASZBD 

Submitted by: 

Atlantic Research Corporation 

Propulsion Division 

751 1 Wellington Road 

Gainesville, Virginia 22065- 1699 

TR-PL-12126 3 1 December I986 
FOREWORD 

This Final Report is structured in two volumes. Volume I includes Task I, 

Conceptual Design Studies, and Task 11, Preliminary Development and Verification (DatV) 

Plan. Volume I1 consists of Task 111, Capability Assessment, and Task IV, ROM Cost 

Estimate. Volume I1 contains proprietary data and shall not be disclosed outside the 

Government. 

ii 
TABLE OF CONTENTS 

VOLUME I - FINAL REPORT 

- PAGE 

1 .O INTRODUCTION 

1.1 OBJECTIVE AND APPROACH 

1.2 SUMMARY AND CONCLUSIONS 

2.0 CONCEPTUAL DESIGN STUDY 

2.1 PRELIMINARY DESIGN CONCEPT 

2.2 CASE 

2.3 CASE JOINTS AND SEALS 

2.4 ASBESTOS-FREE INSULATION 

2.5 NOZZLE DESIGN 

2.6 IGNITER 

3.0 PERFORMANCE 

3.1 MOTOR BALLISTIC PERFORMANCE 

3.2 JOINTS AND SEALS 

3.4 NOZZLE DESIGN 

3.5 IGNITER 

3.3 ASBESTOS-FREE INSULATION 

4.0 PRELIMINARY DEVELOPMENT AND VERIFICATION PLAN 

4.1 DESCRIPTION 

4.2 OBJECTIVE AND SUMMARY 

4.3 MANAGEMENT 

4.4 SRM DESIGN AND ANALYSIS 

4.5 COMPONENT AND SUBSCALE TESTS 

4.6 COMPONENT AND MOTOR MANUFACTURING 

4.7 SYSTEM LEVEL TESTS 

4.8 FACILITIES, EQUIPMENT, AND TRANSPORTATION 

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1 .o INTRODUCTION 

National Aeronautics and Space Administration's (NASA's) George C. Marshall 

Space Flight Center selected Atlantic Research Corporation's (ARC'S) Propulsion Division to perform conceptual design studies of a Block I1 Space Shuttle Solid Rocket Motor 

(SRM). The ground rules for this conceptual motor design included the following: 

The existing performance, design, and verification requirements contained 

in Specification Number CPWI-3300, Part 1, for the Space Shuttle High 

Performance Solid Rocket Motor were used as the baseline requirements 

document for proposed design concepts except as changed and/or amplified by NASA. References to specific design characteristics such as 

segmented cases and other motor design characteristics and/or specif ication requirements unique to Morton Thiokol Incorporated's design were 

deleted; in those instances, ARC selected the specific design characteristics that best suited our overall motor design concept. 

Design concepts essentially duplicate the outside geometry of the current 

Space Shuttle SRM and its interfaces with other Space Shuttle elements so 

that impact to the aerodynamic and dynamic characteristics of the Space 

Shuttle vehicle is minimized. 

Design concepts do not use asbestos-f illed insulation materials. 

Design concepts are not constrained to the current propellant formulation, 

but rather they provide the capability to successfully perform over an 

equivalent polybutadiene acrylonitrile acrylic acid terpolymer (PBAN) 

propellant formulation burn rate range of 0.360 to 0.400 inch/second. For 

information, the performance data contained in CPW 1-3300 is based on a 

PBAN propellant formulation with a target burn rate of 0.368 inch/second 

at 60'F. 

Performance is in accordance with the requirements contained in 

CPWI-3300 from sea level to 200,000 feet over a propellant mean bulk 

temperature (PMBT) range of +40° to +90°F after being subjected to the 

natural and induced environments specified in paragraphs 3.2.7. I and 

3.2.7.2, respectively, of CPWI-3300 to the extent that the PMBT range of 

+40° to +90°F is not violated. 

Any Criticality I, IR, 2, and 2R pressure seal is fully redundant and verifiable (inspected and leak tested in the prescribed functional location). 

Further, no seal requires pressure actuation to perform its designed function. 

1- 1 
Verification methods prescribed in Section 4.0, Table V, of CPW 1-3300 

were not used; the method of verification is a product of this study contract. 

Appendix 10, deviations, and any references to approved deviations were 

deleted from CPW 1-3300. 

1.1 06JECTIVE AND APPROACH 

The objective of the Block I1 SRM Study Program was to define a conceptual 

design that offered 

Equal or greater performance, 

Equal or lower cost, 

Improved flight safety, reliability, and confidence, 

Compatibility, with the current Space Shuttle vehicle and launch facility. 

The contract study flow diagram shown in Figure 1.1.1 schematically 

describes the approach chosen to satisfy the program objectives over the period of 

September 3, 1986 to December 30, 1986. The four interrelated work tasks were performed on schedule in the sequence shown. 

The following design options and trade studies were defined at the beginning 

of the program: 

Segmented versus monolithic design, 

Case, 

Joint and seals, 

Insulation, 

Nozzle, 

Igniter. 

The following evaluation criteria were also defined at program start: 

reliability, 65 percent; cost, 20 percent; and performance, 15 percent. 

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1.2 SUMMARY AND CONCLUSIONS 

ARC evaluated segmented and monolithic SRM design concepts with emphasis 

on joints and seals. Particular attention was directed to eliminating deficiencies in the 

SRM High Performance Motor (HPM) identified in the Report of the Presidential Commission on the Space Shuttle Challenger Accident, during the August 14, 1986 SRM 

Briefing to Industry, and through detailed review of SRM literature, 

The selected conceptual design is described and discussed in Section 2.0. 

ARC'S selection of this concept was driven by the following conclusions: 

An in-line bolted flange field joint is technically superior to either the 

original tang and clevis pinned case joint or the improved "capture 

feature" tang and clevis; however, the bolted flange field joint is heavier. 

The selected NASA Langley Research Center (LaRC) joint offers minimum joint sealing surface deflection. 

The heavier bolted joint weight penalty may be negated by extending the 

current case length to equal SRM casting segment length. ARC believes 

that case founding, forging, and heat treatment facility investments are 

cost effective. 

1-4 
2.0 CONCEPTUAL DESIGN STUDY 

This section describes and documents Atlantic Research Corporation's recommended design concept for a Block I1 Space Shuttle SRM. 

Trade studies were conducted to select specific SRM design features and 

materials in the following areas: 

Motor case, 

Joints and seals, 

Asbestos-free insulation, 

Propellant and liner, 

Igniter, 

Nozzle. 

Design approach (segmented vs. monolithic), 

These trade studies were documented in the Mid-Term Study Report, which is 

included as Appendix A of this report. The detailed trade studies and analyses that led to 

the recommended design concept documented herein may, therefore, be found in 

Appendix A. 

2-1 
2.1 PRELIMINARY DESIGN CONCEPT 

The recommended Block I1 SRM preliminary design concept is depicted in 

Figures 2.1.1 and 2.1.2 and is described below. 

The SRM is a segmented design having casting segment lengths identical to 

those of the Space Shuttle Mission 51-L design. However, longer D6AC case segments 

are utilized, thereby eliminating all but two of the 51-L type factory joints. The three 

field joints connecting the four casting segments are of the inline bolted flange type, 

each fastened with high-strength steel studs with Inconel 718 nuts on each end. Each 

field joint incorporates redundant metallic O-ring face seals. 

The nozzle-to-case joint also incorporates redundant metallic face-type 

seals. In addition, all internal insulation joints are of the unvented type with a labyrinth 

path that precludes direct exposure of the joint seals to hot combustion gases. Mating 

insulation joints are filled with low-strength, high-strain room temperature cure sealant. Stress relief features are incorporated in the insulation near the mating joints to 

permit relative motion of the insulation components without overstressing the insulation 

joint sealant. 

The propellant formulation and grain configuration of each casting segment is 

identical to the High Performance SRM design. 

The case insulation design is a hybrid system to optimize weight and performance. A Kevlar/silica/Hypalon material is used next to the case wall because its low 

thermal diffusivity provides the optimum thermal protection for the reusable case. To 

provide erosion protection near field joints and in areas that are exposed during propellant burn such as the aft case, the Hypalon insulation will transition to an NBR/phenolic 

with boric acid filler (USR-3800). The molded inhibitors will also be made from 

USR-3800. The castable liner will be a carboxyl-terminated polybutadiene (CTPB) material for compatibility similar to the current liner material with the asbestos fibers replaced with another filler material. 

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The nozzle configuration is basically the same as the 51-L configuration 

except that certain materials have been changed to eliminate asbestos and/or to eliminate pocket erosion problems. Also, internal joints have been reconfigured as needed to 

provide redundant seals. 

The preferred igniter design consists of an integral igniter adapter and case 

with a bolt-on aft closure formed from 200 maraging steel. The igniter assembly is 

insulated with Kevlar and silica-filled Hypalon and loaded with 18 percent aluminized 

HTPB propellant. All joints are sealed using T-ring variants and metal C-rings. 

2-5 
2.2 CASE 

The motor case trade study, as shown in Appendix A, resulted in the selection 

of a double-length (casting segment) case fabricated from D6AC and using the roll forging method of metalworking. The selection of this configuration has been substantiated 

by a review of the concept with Ladish Co., Inc. to verify the technical capability to 

produce this configuration and to assess the cost impact of a double-length case segment. 

This assessment compares the technical challenge of a double-length case section with 

that of a single segment case. The newly selected joint configuration was included as a 

consideration in technical and cost evaluation. The motor case design for a typical 

center casting section is shown in Figure 2.2.1, which is included as Attachment I of this 

report. 

Ladish has reviewed thoroughly the double-length motor case, which incorporates the bolted flange design. The configuration can be produced by using the following forge 

machining: 

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sequence, which incorporates the final heat treatment prior to final 

Receive stock, 

Heat, 

Press upset, 

Edge roll, 

Clean pre-form, 

Reheat, 

Extrude and pierce, 

Clean donut, 

Condition donut, 

Reheat, 

First ring roll operation, 

Clean ring, 

Condition ring, 

Reheat, 

Second ring roll operation, 

Hot size, 

Heat treat and second size, 

Machine pre-form blank, 

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Ultrasonic test, 

Roll reduce, 

Stress relieve, 

Cold size diameter, 

Rough machine clevis ends, 

Ultrasonic test, 

Final heat treat in a fixture, 

Ship in a fixture. 

The following is a discussion of the above procedure and is intended to clarify 

the operations. Comments are provided on the forging units that will be required to 

complete the operation successfully, Section 2.3, Volume 11, will address the cost of 

capital equipment required to produce the required length cylinder along with the increased thickness necessary to obtain the bolted clevis joint. 

Ladish will purchase 5O-inch-diameter, 49,500-pound billets (see Figure 2.2.2) 

of Ladish D6AC material from LTV for this requirement. This size is considerably larger 

than the present 40-inch-diameter, 3 1,000-pound billets used to make the single-length 

case. This increased site and weight is beyond LTV's present melt capabilities and will 

require new facilities for melting and handling. LTV is confident that it can maintain the 

present quality level in the larger stock size, but LTV will require a tryout heat to verify 

quality and refine its vacuum-arc remelting operation. 

The next operation is designed to prepare the stock for making the forged 

donut (see Figure 2.2.31, This operation can be performed on equipment existing at 

Ladish. This operation represents low technical risk to motor case fabrication technology. The intermediate cleaning and conditioning operations are routine and will not 

require any additions or modifications to equipment. 

The donut configuration, shown in Figure 2.2.4, is dictated by the roll reduced 

pre-form blank. Since the bolted flange design is thicker than the present clevis, the 

proportions between the cylinder section and the clevis becomes exaggerated in the 

donut. The tonnage of the present press is limited to finished flange thicknesses below 

3.100 inches. 

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The hot roll operations (see Figures 2.2.5 and 2.2.6) can be produced using 

Ladish's current equipment; however, due to length and configuration, new ring roll tools 

will be needed. No major technical risks are expected with this change. 

The hot size operation (see Sketch #6 in Appendix B of Volume 11) is low risk 

and can be performed on equipment existing at Ladish. However, new sizing segments 

will be required due to the longer length pre-form size. 

Machining of the pre-form blank (see Sketch 416 in Appendix B of Volume 11) 

from the rolled ring should present no technical challenge. No new equipment will be 

needed for this operation. 

Ladish has developed a method of using immersion sonic testing the large preforms that will be used. Slight modification of Ladish's equipment to accommodate the 

longer length is anticipated. 

The next operation is the flow turning of the blank to its total length. A new 

machine will be required to manipulate the longer length and the starting pre-form 

blank's thicker wall section. The new machine will be patterned after Ladish's successful 

flow turn machine. No major problems in using this new piece of equipment are anticipated. 

After flow turning, all handling is accomplished using fixtures to prevent 

damage to the ring due to its long length and thin wall section. 

A new furnace will be required for the stress relieving operation. The case 

can be hung vertically in the furnace and fixtured to keep it round. This operation varies 

from present practice, since the 160-inch cylinders fit into existing furnaces and do not 

require the extensive fixturing necessary for the longer length. This operation is the 

first in this production sequence that varies to any great extent from the present practice. 

The cold sizing operation will require a new press because of the added 

length. The press will be designed with added segment travel to accommodate the bolted 

flange joint, which has the added stock on the inside diameter. New sizing tools will be 

needed for the added length. This operation will be essentially the same as the current 

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sizing practice. The machining of the bolted flange joint will be done on the current 

equipment modified for the increased length. Again, this operation is similar to existing 

practices with only the machine contour changing. 

The final ultrasonic test of the roll formed cylinder will use the current 

inspection method but will require new facilities due to the increased length and the 

requirement to fixture the ring during sonic testing. This added weight will require 

stronger rollers and tank. Sonic testing these pieces in a fixture should not present any 

new technical challenges. 

The final heat treatment (see Sketch #6 in Appendix B of Volume 11) will take 

place in a new furnace. This operation is the only addition to the current process. A new 

furnace will be designed and installed for this operation. Again, fixturing will be required to maintain roundness. The heat treatment facility will include salt bath quench 

tanks. 

Presently, components are shipped on flat cars blocked into the vertical 

position. The length of the double-length case will not allow this shipping configuration. 

The new motor cases will be shipped in the horizontal position with an internal fixture on 

rail cars that incorporate a shipping cradle. This method should provide sufficient transportation protection and maintain roundness during transportation. 

The procedures that will be used to produce the double-length motor case 

incorporating the bolted flange are basically the same as those presently used for the 

160-inch long motor case. The added length provides some problems that are not encountered with the present cylinder. The first is the new starting billet size and weight. 

Historically, the startup of new melt facilities requires debugging and introduces unknowns into the steel quality. Process characterization is, therefore, planned early in 

the Development Phase. 

The second problem is handling of double-length motor cases. Ladish plans to 

develop new handling equipment and fixtures. Also, a learning process is planned during 

initial process development. Ladish is confident that after these two issues are understood, the process will be technically sound and cylinders can be made to the required 

quality and dimension requirements. 

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2.3 CASE JOINTS AND SEALS 

This section discusses the structural analyses and the salient structural features of the metal case field joints and nozzle joint and of the case and nozzle insulation 

joints. Additionally, rationale for the selection of the seals is given and a discussion of 

their characteristics and background usage is presented. 

2.3. I METALLIC CASE JOINT 

2.3.1.1 JOINT SELECTION 

The selection process for the in-line bolted field joint is given in detail in the 

Mid-Term Report of Appendix A. In brief, the selection was driven by safety and teliability considerations, including the following: 

The in-line bolted joint allows a face seal arrangement. This arrangement, in turn, allows the use of metallic seals, which are known to be 

much more temperature resistant than rubber O-rings (2,500°F versus 

500°F). They are also more resilient than rubber and will track instantaneously any flange opening. Furthermore, the metal is not sensitive to 

cold ambient temperatures. 

The bolted joint is amenable to analysis via three-dimensional finite 

element techniques, including the effect of pre-loading the bolts. 

Because of the above, the NASA Langley Research Center was able to 

optimize the bolt circle position relative to the membrane shell middle 

surface and, hence, minimize flange opening at the seal positions. These 

dimensions are incorporated into the current design. 

2.3.1.2 CONFIGURATION AND STRUCTURAL FEATURES 

Figure 2.3.1 shows the in-line joint that was analyzed and its critical dimensions. The design is identical to that of the November 23, 1986 LARC design with the 

following exceptions: 

The ARC design has an Inconel O-ring for both the primary seal and the 

secondary (redundant) seal. The seals are discussed in detail in Section 

2.3.5. 

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The ARC design, as analyzed herein, utilizes a shear lip around the 

inboard circumference of the lower segment flange. This shear lip is 

used primarily to make rounding and fit-up easier during field assembly. 

This design has been modified slightly to thin the flange from 3.5 inches 

to 3.1 inches (inside to outside). 

Bolting consists of 150 1.1 25-inch-diameter, high-strength studs from SPS, 

type MP 150. This scheme was found by LaRC to be reasonably efficient with respect to 

spacing, metal O-ring size, and flange thickness. Alcoves are large enough to allow the 

use of hydraulic tightening of the special nuts while employing an ultrasonic transducer 

for the LaRC pulsed phase locked loop system, which will be used to measure accurately 

stud pre-tension.") The transducer is affixed magnetically to the upper end of the 

stud. Four systems will be used simultaneously during field assembly. 

Finally, we note that the case membrane shell is tapered for 7 inches from 

0.479 inch nominal thickness to 0.90 inch nominal thickness at the alcove "roofstf, thus 

minimizing the effects of the discontinuity. 

2.3.1.3 DESIGN CONDITIONS 

Table 2.3.1 outlines the major structural requirements for the SRM redesign. 

We note here that the ultimate safety factor of 1.4 was applied to both the equivalent 

axial load of 18.26 x 10 pounds and the internal MEOP of 1,004 psi in the stress analysis. This safety factor was applied to ensure that the flange gap opening would meet 

the required safety factor, as well as the stressed metal parts. 

6 

Allison and Heyman, "Nondestructive Ultrasonic Measurement 

of Bolt Preload Using the Pulsed Phase Locked Loop 

Interferometer," The Second Symposium on Welding, Bonding, 

and Fastening, NASA Langley Research Center, October 23 - 25, 

1984. 

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TABLE 2.3.1. JOINT AND SEAL DESIGN CONDITIONS. 

1. MEOP = 1,004 PSIG. 

2. ULTIMATE SAFETY FACTOR = 1.4. 

3. YIELD SAFETY FACTOR = 1.2. 

4. MAXIMUM MOMENT AT UPPER FIELD JOINT, M = 68 x lo6 IN-LB. 

5. EQUIVALENT LOAD: 

Weq=PnR 2 +- 2M 

R 

AT R = 72.0", P = 1,004 PSIG 

= 18.26 X 10 6LB WEQ 

6. MATERIAL: 

D6AC CASE 

CASE ULTIMATE = 200 KSI 

CASE YIELD = 185 KSI 

BOLT/STUD MULTIPHASE ALLOY (MP 150) 

ULTIMATE = 260 KSI. 

7. MAXIMUM ALLOWABLE FLANGE OPENING AT SEALS = 0.012 IN. 

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2.3.1.4 BOLT STRUCTURAL ANALYSIS 

The ultimate bolt load is as follows: 

= 1.4 (18.26 no6)/150 

= 170,400 pounds/stud 

FU 

Pre-tensioning of the studs is held to 70 percent of their ultimate strength, or 

= 0.70 (234,000) 

= 164,000 pounds 

FP 

Thus, the pre-tension will be lost as the ultimate design load is approached; however, no 

problem will occur as long as the flange opening is kept to a safe level. The subsequent 

finite element analysis shows this to be the case. However, insofar as the studs are 

concerned, the ultimate safety factor (SF) becomes (with a listed strength of 234,000 

pounds/bolt): 

- 234,000 (1.4) 

SF" - 170,400 

= 1.92. 

This safety factor provides a margin over and above that which is required. 

2.3.1.5 CASE WALL THICKNESS 

Although the membrane wall portion of the motor case is presumably unchanged from the original lightweight HPM design, a few calculations are given regarding 

the membrane stress state and the safety factors to provide continuity with the overall 

joint analysis. The mean shell radius is taken as 72.8 inches and the minimum wal1,as 

0.459 inch for these calculations. 

BIAXiAL EFFECTS 

It is important to account for the affect on ultimate strength that the twoto-one stress field found in a thin-walled pressure vessel might have. For the class of 

low alloy steels, similar to D6AC, the MIL-HDBK-5D suggests a factor of 0.90 to be used 

2-19 
to calculate strength enhancement in a two-to-one biaxial stress field. 

effective stress, ue, to be compared to the uniaxial tensile strength (u" = 200 ksi) 

is given by u = 0.9 ue, where ue is the induced hoop stress. 

That is, the 

e 

SAFETY FACTOR 

At the MEOP, 

1,004 (72.81 

0.459 u= e 

= 159,240 psi. 

Hence, 

u = 0.9 a@ e 

= 143,300 psi, 

and the ultimate safety factor is 

200,000 

SF = 143,300 

= 1.4. 

Thus, the design requirement is met and case segments with the minimum wall at 0.459 

inch can, in fact, be used as forward segments where the MEOP is taken as 1,004 psi. 

We note that the enhancement factor for the classical von Mises stress is 

0.866 for a two-to-one biaxial stress field, such that the von Mises failure criterion 

reasonably represents the D6AC material. For this reason, it was decided to present the 

NASTRAN stress contours (Section 3.2. I .7) in the form of von Mises effective stress. 

MODEL SHELL STRESS 

The NASTRAN shell von Mises stress away from the joint should be 

* (0.866) (1.4) (1,004) (72.8) 

0.479 u= 

= 185,000 psi. 

2-20 
This value agrees very well with the finite element results (Figures 2.3.6 and 2.3.7). 

2.3. I .6 FIMTE ELEMENT ANALYSIS 

A PATRAN/NASTRAN finite element model of the case, joint, and bolts was 

constructed as shown in Figure 2.3.2. As shown, the model is a 1/300 circumferentially 

repeating segment of the full-up motor. The axial length of the model is 56 inches, 

which is sufficient to capture the membrane behavior of the case away from the joint. 

The model consists of 1,724 solid elements. The stud and nuts are connected to the 

flanges only by linear contact elements, as are the mating flange surfaces. 

The model is loaded according to Table 2.3.1 with the 1.4 ultimate factor 

included in the axial end load and in the internal pressure load. These "ultimate" loadings 

are reflected in all stress and displacement results. The applied pretension of the studs 

was induced into the NASTRAN solution as a cooling of the stud to an equivalent a A T 

of -0.0090. The stud pre-tension was 58 percent of the ultimate strength, or 135,000 

pounds. This value is lower than the intended value of 164 ksi, but it should have no 

significant effect on the calculated state of stress in the case or joint. Displacement 

values (flange opening) might be slightly greater than those with the full pre-tension. 

2.3.1.7 FINITE ELEMENT RESULTS 

DEFORMATIONS 

Figure 2.3.3 shows the deformation pattern of the bolted joint in the vicinity 

of the flanges. The "heel-toe" action is clearly seen with contact indicated over most of 

the area of the seal grooves. Figure 2.3.4 shows the overall joint and shell deformation 

patterns. These patterns indicate that the model was, in fact, long enough (56 inches) to 

capture the membrane shell behavior away from the joint discontinuity. Figure 2.3.5 is a 

plot of the relative displacement between flanges (gap opening). The profile of the 

O-ring grooves is shown for reference. Only at the secondary seal under the alcove's 

gusset plate (halfway between bolts) is there a positive separation, and this value is a 

maximum of 0.0031 inch over the backup O-ring. With a 0.012-inch O-ring recovery 

capability, the ultimate safety factor on deflection (gap opening) is estimated to be 

(1.4) = 5.4 0.012 

SF = 0.0031 

2-21 
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OF POOR QUALITY 

2-22 

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FIGURE 2.3.3. DEFORMED GRID AT FLANGES. 

2-23 
ORIGINAL PAGE IS 

OF POOR QUALITY: 

.... l.1’ ... 

F IGURE 2.3.4. OVERALL CASE AND JOINT DEFORMAT ION. 

2-24 
ORIGINAL PAGE 1s 

OF POOR QULITy 

. 

DISTANCE ACROSS FLANGE 

FIGURE 2.3.5. FLANGE FACE SEPARATION. 

2-25 
STRESS RESULTS 

Figures 2.3.6 and 2.3.7 show stress contour plots on isometric views of the 

lower and upper segments. The color patches cover the indicated ranges of the values of 

the von Mises stress. These values agree well with the von Mises membrane stress in the 

shell (away from the joint) of 185 ksi as previously calculated in Section 2.3.1.5. Stresses 

throughout the tapered shell walls and the alcove gussets and roofs are satisfactorily 

low. However, stress concentration regions appear at the edges of the bolt holes on the 

respective mating surfaces of the upper and lower flanges. These areas of high stress are 

due to the presence of the holes in what would otherwise be a fairly uniform hoop tension 

field. Figures 2.3.8 through 2.3.11 show the von Mises stress contours. The very local 

nature of the stress concentrations is apparent in that the yield regions do not penetrate 

to the alcove sides of the flanges. The material that is stressed above about 185 ksi may, 

of course, yield, but it will not in any way threaten the ultimate strength of the joint 

structure. The maximum strain at the edge of the bolt holes can be estimated by 

261,500 

29 x lo6 

E= 

= 0.0090 inch/inch 

With a minimum elongation of 6 percent for the D6AC material, the ultimate safety 

factor would be 

= 9.3. 

In a final Design Phase, this condition would be more rigorously analyzed via a nonlinear 

plasticity version of the NASTRAN code. We also note that the results shown herein are 

for an ultimate load condition. At the MEOP condition, the peak stress concentration 

reduces to 

261 500 

O=1.4 

= 186,800 psi. 

This pressure would cause, at worst, a tiny spot of yielded material. 

2-26 
ORIGINAL PRGE 

COLOR PHOTOGRAPH 

FIGURE 2.3.6. STRESS CONTOURS ON 

LOWER SEGMENT. 
ORIGINAL PAGE 

COLOR PHOTOGRAPH 

. . ..... .... :::( 

FIGURE 2.3.7. STRESS CONTOURS ON 

UPPER SEGMENT. 
ORIGINAL PAGE 

COLOR PHOTOGRAPH 

2- 

I 

I FIGURE 2.3.8. STRESS CONTOURS ON FLANGE - MATING 

SURFACE - UPPER SEGMENT. 
ORIGINAL PAGE 

COLOR PHOTOGRAPH 

.... .-- 

.... .- -1 

.- * -d *. , 

* ... I, C. *.. ,-. I 1- 

- 

FIGURE 2.3.9. STRESS CONTOURS ON FLANGE - VIEWED FROM 

_. 

SI * 

7,’ 

... .... -- 

j 

J-.- ALCOVE - UPPER SEGMENT. 
ORIGINAL PAGE 

=OR PHOTOGRAPH 

- .. . . . . . . . . . . -~ ~. . . . . 

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-. r 

I. 

FIGURE 2.3.10. STRESS CONTOURS ON FLANGE - MATING 

SURFACE - LOWER SEGMENT. -_ -- 
............ 

............ 

............ ............ 

............. ............ 

............ 4 ............ 3 

$ ............ s 

I' ............ i ............ i ........... 

.............. 

............ 

............ 

............ f 

ORIGINAL PAGE 

COLOR PHOTOGRAPH 

FIGURE 2.3.11. STRESS CONTOURS ON FLANGE - VIELIED .. II 1.1 .. FRON ALCOVE - LOWER SEGMENT. 
2.3.1.8 SUMMARY AND CONCLUSIONS 

Important results of the case joint structural analysis are listed below. 

The maximum flange opening (at the ultimate load) over the secondary 

metal O-ring is 3.1 mils. The ultimate factor of safety is 5.4. 

The membrane case safety factor is not affected by the joint design. For 

the "lightweight" case with a 0,459-inch minimum wall and 200,000-psi 

ultimate tensile strength, the ultimate safety factor is 1.4. 

The bolt ultimate safety factor is 1.92. 

Alcove gussets, backwalls, and roofs are relatively lowly stressed. The 

maximum gusset stress is about 170 ksi (at 1.4 x MEOP), and the maximum ultimate safety factor is in excess of 1.65. 

Stress concentrations appear at the edges of the bolt holes. The peak 

stress just reaches the yield strength at MEOP. The high stress regions 

are very localized, and the ultimate safety factor is estimated to be 9.3 

based on plastic strain. 

The metal joint weight over that of an equivalent length membrane shell 

is 1,800 pounds. 

The major conclusion is that the metallic components of the in-line bolted 

joint meet all of the design goals and safety requirements for the HPM baseline operating 

and reuse conditions. Further options were investigated that reduced the width of flange 

and, consequently, reduced the required forging thickness. The structural impact of 

narrowing the flange to 3.1 inches is estimated to be negligible. 

2-33 
2.3.2 CASE INSULATION FIELD JOINTS 

In the first phase of this study, documented in the Mid-Term Report (Appendix A), trade studies and evaluations of various case and insulation joints were performed 

to select baseline joint and seal designs. An in-line bolted flange using two metallic 

O-ring seals were selected as the optimal case joint design. These seals are fixed in a 

compressed state and do not require gas pressurization to actuate sealing. For this type 

of sealing system, an unvented insulation joint was determined to provide the highest 

integrity and reliability. 

The selected case insulation joint (Figure 2.3.12) is an unvented labyrinthtype that utilizes an elastomeric, open channel, stress relief component vulcanized 

within the case insulation. Radially oriented oval channels provide stress relief while 

preventing circumferential gas flow. A low modulus NBR/silica is used for the stress 

relief component to provide high elongation capability. NBR/silica is also used for the 

male portion of the overlap joint to provide a nominal 0.070 inch of compression at static 

conditions. A low-strength, high-strain, ambient-cured RTV adhesive/sealant fills the 

gap inboard of the compressed overlap. Detailed analyses of the insulation joint are 

presented below. 

ARC performed a case insulation structural analysis. The objectives of this 

analysis were twofold: first, the nonvented configuration was analyzed to determine the 

effectiveness of the stress relief component, specifically, the normal and shear stress 

across the adhesive joint; and, second, the system was analyzed with an unbonded joint to 

determine the amount of opening at the unbonded interface. The latter result was necessary to proceed with the gas dynamic and heat transfer analysis of the metal case joint. 

Figure 2.3.13 shows the axisymmetric TEXCAP finite element approximation of the 

grain, insulation, and motor case. The model was extended for 28 inches forward and aft 

of the joint, and a condition of plane strain was assumed for the "cut" ends of the grain. 

An internal pressure of 1,004 psi was used in conjunction with a a AT of 0.0024. The 

latter load assumes assembly at 80°F and a launch bulk temperature of 40°F. 

Stress results are shown in Figure 2.3.14. These results are considered to be 

approximate since a complete laboratory tensile characterization of the viscoelastic 

grain and insulation materials would be required to better pinpoint their respective 

2-34 
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2-35 
FIGURE 2.3.13. AXISYMMETRIC FINITE ELEMENT MODEL - 

CASE JOINT, INSULATION, AND GRAIN. 
MOTOR CASE 

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SHEAR/NORMAL STRESSES 

FIGURE 2.3.14. STRESSES ACROSS INSULATION BONDLINE. 

2-37 
moduli. The stress values presented are considered to be reasonably low and within the 

capability of adhesive systems. All stress normal to the bondline is compressive. 

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Deflection results are shown in Figure 2.3.15. These results were used in the 

thermal analysis below. 

Thermal analyses of the case insulation joint were performed to evaluate the 

following three areas: 

Erosion and charring in the joint region, 

Heating in the stress relief channels, 

Heating in the insulation joint. 

Details of the analytical techniques are contained in Section 2.4. 

Heating rates are fairly benign in the joint region due to protection by the 

forward inhibitor. Convective film coefficients are approximately 30 percent of the 

values along the sidewall, and luminous thermal radiation from combustion gas is negligible due to the close proximity of the inhibitors. In-depth temperatures, erosion rates, 

and heat-affected rates are shown in Figure 2.3.16. The insulation joint is sized such 

that the compressed face of the overlap joint is not thermally degraded during motor 

operat ion. 

Heating rates in the stress relief channels are solely dependent upon local gas 

flow rates in the channel. Thermal evaluation of gap heating rates requires coupling of a 

one-dimensional gas flow module to a two-dimensional conduction network module. 

Positive pressurization rates force hot gases into the channels; the gases transfer heat to 

the surrounding wall and contract, thereby decreasing pressure and causing a low level 

influx of additional hot gases. Conversely, a pressure decrease results in an efflux of gas 

from the channels. Circumferential flow is prevented by the insulation web between 

channels. The HPM pressure profile is fairly regressive, which tends to produce the 

majority of heating during the ignition phase. Oscillatory pressure fluctuations, such as 

those caused by vortex shedding from the inhibitors, and mass transfer effects from 

insulation decomposition were not included in the channel heating rate analyses summarized below. 

2-38 
I 

FIGURE 2.3.15. INSULATION SEAL JOINT OPENING CHARACTERISTICS - 

ADHESIVE ASSUMED TO BE FAILED. 

bU2 = .0509 

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Figure 2.3.17 presents gas temperatures and heating rates along the length of 

the stress relief channel. Gas temperatures are shown to decay rapidly with time from 

ignition and with distance from the channel opening. These heating rates were used in 

heat and mass transfer analyses to determine erosion and degradation of the stress relief 

wall. Results shown in Figure 2.3.18 verify that the web between channels will not erode 

during motor operation, thereby preventing the occurrence of circumferential flow. 

The stress relief component is configured to preclude opening of the unvented 

insulation joint during motor operation. Nevertheless, the joint must be capable of 

operating successfully in a vented mode to ensure that adequate reliability is achieved. 

In a vented mode, the primary heating is due to a combination of radial flow during the 

initial portion of burn and circumferential flow during the remainder of operation. A 

steady 0.5-psi pressure drop was used to determine mass flow rates in the joint. Erosion 

and heat-affected rates in the gap are presented in Figure 2.3.19. Erosion rates are 

sufficiently small to preclude failure of the compressed face of the joint. 

The above analyses verify the thermal and structural integrity of the insulation joint design. Stress relief minimizes both joint stresses and deflections. Erosion is 

sufficiently small to preclude consumption of the web between stress relief channels and 

degradation of the compressed face of the joint overlap. Furthermore, the insulation 

joint is shown to be capable of withstanding circumferential flow if the joint becomes 

vented. 

2-41 
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2.3.3 NOZZLE-TO-CASE METAL JOINT 

2.3.3. I JOINT SELECTION 

The selection process for the nozzle-to-case joint hardware configuration is 

described in Appendix A, the Mid-Term Report, The preliminary selection given in that 

report, the dual face seal approach, was confirmed by subsequent work. In brief, the 

selection was driven by safety and reliability considerations. Some of these considerations are 

The dual face seal joint allows the use of metallic seals, which are much 

more resistant to high temperatures than the rubber O-rings required by 

any design utilizing bore seals. The metallic seals will also track gap 

opening instantaneously, even at cold ambient temperatures, which significantly reduce the tracking capability of rubber O-rings. 

The dual face seal arrangement allows both the primary and secondary 

seals to be positively pre-loaded with the existing attachment bolts. 

In the event of leakage past the unvented insulation joint, the face seals 

are protected from direct impingement of a hot gas jet. 

2.3.3.2 STRUCTURAL DESCRIPTION 

The ARC Block I1 nozzle-to-case hardware joint design, which is shown in 

Figure 2.3.20, incorporates face seals as both the primary and secondary seals. Changes 

required from the 51-1 design include increasing the width of the flange mating surfaces 

to accommodate the revised seals, increasing the nozzle flange thickness to reduce gap 

opening of the primary seal, and the addition of a shear lip on the nozzle flange with a 

radial interference fit. The shear lip eliminates relative radial sliding of the dome and 

nozzle flanges and reacts the resulting radial force in direct shear rather than by friction 

or shear in the bolt. Membrane thicknesses in the case dome and the nozzle fixed 

housing are unchanged from the 51-L design. Critical dimensions are shown in 

Figure 2.3.21. These revisions result in a 740-pound weight impact. 

Also included in the Block 11 design are leak check ports, which allow verification of both the primary and secondary seals in the direction of operation. A rubber 

O-ring bore seal is retained to act as a leak check seal for the primary face seal. 

2-45 
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The nozzle is attached using 100 1.375-LZUNJF bolts, which are pre-loaded to 

6 a total of 16.5 x 10 pounds minimum, approximately 3.9 times the nozzle blow-off load. 

2.3.3.3 DESIGN CONDITIONS 

Table 2.3.2 lists the primary structural requirements used for the Block I1 

joint design. Maximum pressure and thrust loadings were imposed to ensure the structural integrity of the nozzle fixed housing, the motor aft dome, and the nozzle attachment 

bolts. Seal deflection requirements were interpreted to mean that gap openings due to 

the application of limit loads must be kept to less than one-half the seal capability. 

2.3.3.4 FINITE ELEMENT ANALYSIS 

A PATRAN/NASTRAN three-dimensional finite element model (FEM) of the 

nozzle-to-case joint was generated for structural analysis and is shown in Figure 2.3.22. 

The model represents a one-hundredth (3.6 degrees) repeating segment of the motor 

circumference and utilizes symmetry boundary conditions to simulate the cyclic nature 

of the structure. The entire length of the nozzle fixed housing and the motor aft dome 

up to the tangency point were included. This causes the joint area to be sufficiently 

remote from the axial boundaries so that joint deformations are not significantly 

affected by rotation and radial deflection boundary conditions of the forward end of the 

FEM. The attachment bolt was rigidly attached to the motor dome, simulating the 

threaded condition. The nozzle and motor flanges, and the bolt head and nozzle flange, 

were connected using only linear contact elements. A radial shear connection was 

included between the nozzle and motor dome at the shear lip, while no shear connection 

was included between bolt and nozzle, Joint pre-load was simulated by imposing a reduced temperature (-859'F) on the bolt shank, which was given a coefficient of thermal 

expansion of 6.6 x in-in/OF. The resulting pre-load was 154,200 pounds, which was 

only slightly less than the 165,000 pounds that was desired. This small difference did not 

significantly affect the results. 

Limit loads were applied to the NASTRAN model according to Table 2.3.2. 

Pressure was applied along the entire internal surface and in the radial gap between the 

nozzle and the dome attachment ring, down to the primary seal. A nonlinear analysis 

was conducted in which contact loads between parts were determined consistent with the 

relative flange deformations by iteration within the NASTRAN solution procedure. 

2-48 
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i TABLE 2.3.2. JOINT AND SEAL DESIGN CRITERIA. 

JOINT HARDWARE 

1. 

2. 

3. 

4. 

5. 

6. 

MAXIMUM PRESSURE = 909 PSI (LIMIT). 

THRUST = 3.31 x lo6 POUNDS (LIMIT). 

APPLIED LOAD AT FORWARD END OF FIXED HOUSING = 1.16 x lo6 POUNDS (LIMIT). 

ULTIMATE SAFETY FACTOR = 1.4. 

FIELD SAFETY FACTOR = 1.2. 

PRESSURE ACTS UP TO PRIMARY SEAL. 

DEFLECTION AT SEALS (GAPS) 

1. SEALS SHALL DEMONSTRATE SEALING CAPABILITY AT: 

A. 

B. 

TWO TIMES MAXIMUM GAP SIZE AT MAXIMUM EXPECTED TRACKING RATE. 

TWO TIMES TRACKING RATE AT MAXIMUM EXPECTED GAP SIZE. 

2-49 
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The total model deformations for the limit load condition are shown in 

Figure 2.3.23, while Figure 2.3.24 shows deformations in the joint region. In both of 

these computer plots, deflections are multiplied by a scaling factor for display purposes. As can be seen from these figures, the nozzle housing tends to move inward while 

the motor dome tends to deflect outward, which is consistent with the relative pressure 

loadings. This mismatch in radial deformations is one of the primary causes for the 

opening of the joint. The other major cause is the eccentricity between the line of 

action of the blow-off load carried down through the fixed housing and the reaction at 

the bolted interface. This combination of loads leads to gap opening at the nozzle-todome interface as shown in Figure 2.3.25 for the limit load condition. This is a view 

looking aft at the mating surface, and the numbers shown indicate the axial gap opening 

at each location. These results show that the gaps are uniform in the circumferential 

direction with essentially no waviness. This deflection pattern also indicates that the 

flanges remain in intimate contact outboard of a radius, which is just inboard of the bolt 

center line. 

Computer-generated contour plots were used to evaluate the maximum 

stresses that were used in margin of safety calculations. Figures 2.3.26 through 2.3.30 

illustrate these stress distributions in critical regions for the limit load condition. Shown 

in Figures 2.3.26 and 2.3.27 are von Mises stress distributions in the forward and aft 

faces of the nozzle flange, respectively. Cross-sections of the nozzle housing and the 

motor dome, with the von Mises stress contours shown, are included in Figures 2.3.28 and 

2.3.3.29, respectively. Finally, Figure 2.3.30 is the axial stress distribution in a bolt 

cross-section. A summary of the stress state in the metal parts is given in Figure 2.3.31, 

where the maximum stresses are noted at critical locations along with the resulting 

margins of safety. Critical margins were all positive, which verifies the structural 

integrity of the joint hardware. 

2.3.3.5 JOINT SEALING CAPABILITY 

The maximum gap openings of the inboard edge of each seal groove were 

determined for limit loads to evaluate the joint sealing capability. These gap openings 

are illustrated in Figure 2.3.32. Also shown in this figure are the seals themselves, 2 

0.375-inch-diameter, silver-plated, Inconel metal O-rings. The seals are described more 

fully in Section 2.3.5. The maximum allowable gap opening (springback) recommended by 

the manufacturer for these Inconel O-rings in this application is 0.012 inch (see 

2-51 
ORIGINAL PAGE 1s 

OF POOR QUALITY. 

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2-61 
Section 2.3.5). As shown in the figure, the maximum gap opening at the primary seal was 

found to be 0.0051 inch, while the corresponding value for the secondary seal is 

0.0017 inch. Since each of these is less than one-half the manufacturer's recommendation, sealing capability at twice the maximum predicted gap is verified. There is some 

conservatism in this evaluation since the actual springback for the seals has been measured to be 0.033 inch for the 20-percent minimum compression used here. 

The gap openings at the nozzle-to-case joint sealing surface are a somewhat 

nonlinear function of motor pressure. Figure 2.3.33 shows the gaps at the primary and 

secondary seals as a function of motor pressure. By using these data in conjunction with 

the ignition pressure transient time history from the DM-1 static firing, the gap openings 

versus time were found as shown in Figure 2.3.34. Note that these are smooth curves 

without the sharp discontinuities (jumps) associated with radial slippage (rounding) for a 

bore seal. The maximum gap opening velocity, shown in Figure 2.3.34 to be approximately 0.029 inch/second, appears to be well within the tracking capability of the metal 

O-rings used. However, since no hard data exists for this seal characteristic, this area 

2.3.3.6 SUMMARY 

In summary 

requires some developmental testing to verify the seal's capability. 

the dual face seal approach was found to be a simple effective 

design solution for reliably sealing the shuttle SRM nozzle-to-case joint. The primary 

drawbacks were found to be weight impact and the inability to use existing parts. However, the design presented herein represents a single design iteration, and we believe that 

further development would yield a design that minimizes these impacts. For example, it 

appears that utilizing larger bolts on a bigger bolt circle would be a more weighteffective approach, and expanding the bolt circle for the current size bolts might allow 

the use of existing domes that did not yet have the bolt holes drilled. Design optimization would be part of the Development Program for this approach. 

2-62 
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FIGURE 2.3.34. JOINT SEAL DYNAMICS. 

2-64 
2.3.4 NOZZLE-TO-CASE INSULATION JOINT 

Trade studies performed during the first phase of this study (documented in 

Appendix A) selected an Inconel C-ring for the inboard seal and an Inconel O-ring for the 

outer seal. Neither seal requires pressure to actuate sealing, since both seals are maintained in a compressed state. An unvented notzle-to-case insulation joint was selected 

as the best configuration for this sealing system. 

The selected nozzle-to-case insulation joint (Figure 2.3.35) is an unvented 

type filled with a high-strain, ambient-cured RTV adhesive/sealant. An open channel, 

NBR/ silica stress relief component is vulcanized into the USR-3800 case insulation. 

Radially oriented oval channels in the stress relief component provide a means for both 

tensile and compressive deformation while minimizing stresses in the joint and preventing circumferential flow. A silicone foam with a high temperature capability, 

bonded along the overlap portion of the joint, provides redundant flow baffling while 

preventing extrusion of the RTV adhesive/sealant into the C-ring groove during assembly. 

Detailed structural and thermal analyses of this joint are presented below. 

Structural analysis of the nozzle-to-case insulation joint was conducted using 

a two-dimensional TEXGAP finite element model, which is shown in Figure 2.3.36. This 

simulation included all the critical features of the joint. The steel structure of the 

Block I1 SRM design nozzle housing and motor aft dome were modeled along with their 

respective insulations: NBR/phenolic over a layer of Hypalon for the motor dome and 

carbon/ phenolic for the nozzle housing. The housing insulation layer of glass/phenolic 

was neglected, with the carbon/phenolic properties used in that region. A 0.100-inchthick bondline was placed between the NBR/phenolic and carbon/phenolic at the joint 

above the step. An open gap was left below the step. The elastomeric stress relief was 

modeled with an orthotropic material having appropriately reduced R-2 properties and 

negligible properties in the hoop direction, simulating the open holed geometry. The 

bolts were modeled by a ring rigidly attached to the motor dome with the correct axial 

stiffness. This was accomplished using an orthotropic material having the appropriate 

axial modulus and negligible hoop modulus. Thin, stiff "contact" elements were included 

under the bolt %head" and between the mating flanges. By iteration, these elements were 

added or removed until a consistent pattern of compression in the contact elements and 

separation, where the contact elements are absent, was achieved. Joint pre-load was 

provided by applying a negative a A T to the bolt to give a net axial pre-load of 16.5 x 

10 pounds. 6 

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The model was used to evaluate stresses, strains, and displacements for two 

distinct loading conditions. The first consisted of limit pressure applied to the interior 

surface of the insulation and in the stress relief slot. For this condition, it was assumed 

that the secondary adhesive bond remained intact and, thus, no pressure was applied in 

the gap between the nozzle and case insulation. The deformed shape of the FEM for this 

load case is shown in Figure 2.3.37. The deflections in this plot have been multiplied by a 

factor of 5.0 for presentation purposes. The analysis showed that the deflection across 

the stress relief is influenced both by the metal joint rotating open and by radial deflection of the nozzle housing and motor dome forward of the joint. The undeformed and 

deformed slot width at the forward end of the stress relief is shown on the figure to 

illustrate deflections in this critical region. The stress/strain state in the insulation joint 

for this load case is summarized in Figure 2.3.38. In most areas of interest, maximum 

strain is the governing failure criteria; therefore, the maximum strain is shown at critical points along with the associated margin of safety. All margins of safety are positive, 

which verifies the structural integrity of the joint in the unvented condition. 

The other loading condition considered was based on a fail-safe approach. It 

was assumed that the secondary adhesive bond had failed, allowing the gap between 

nozzle and dome insulation to open. Therefore, pressure was applied to both faces of the 

gap down to the primary seal location in addition to the previous loads. The deformed 

shape of the FEM for this condition is shown in Figure 2.3.39. The stresses and strains in 

the insulation for this condition are not critical. The original gap width (bondline) and 

the resulting width after deformation are shown on the figure. The latter is utilized in 

the following thermal analysis. 

The joint must be capable of operating successfully, even if for any reason 

the joint becomes vented, to ensure positively that adequate reliability exists under all 

operating conditions. Thermal analyses of the joint in a vented mode were performed for 

this purpose. Thermal analyses were also performed to evaluate heating in the stress 

relief channels. 

Convective heating rates are substantial in the nozzle-to-case joint region 

due to the initial high velocity flow between the grain and insulation and due to apparent 

recirculation in this region upon grain burnback. Flowfield analyses of this region, performed by MTI, were inadequate in predicting the observed insulation material-affected 

rates. A review of the analytical techniques showed the flowfield calculations were 

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based upon potential flow; a solution to the elliptical form of the Navier-Stokes is required to properly account for the flow characteristics in recirculation regions. In the 

absence of meaningful flowfield characteristics and convective film coefficients, the 

thermal analyses presented below were based upon heating rates determined by emperically fitting measured material-affected rates. Figure 2.3.40 shows the predicted erosion pattern in the aft dome region. The stress relief component is shown to have an 

adequate length to function properly in the eroded configuration representing burntime 

conditions. 

Heating in the stress relief channels is modeled by coupling a one-dimensional 

flow module to a two-dimensional conduction solution (see Section 2.3.2). The dominant 

mode of heating is due to convection from a high velocity influx of gas into the channel 

such as occurs at ignition. Rapid pressure fluctuations due to vortex shedding from the 

inhibitors or from rapid changes in recirculation patterns induced by grain burnback were 

not included in the analyses. The results presented in Figure 2.3.41 show that the channel web erosion is dominated by erosion of the surrounding insulation. 

Analytical assessment of the erosion in the joint was based upon venting 

occurring at ignition. Heating is due to a radial flow of gases at ignition followed by 

circumferential flow for the remainder of motor operation. A steady 0.5-psi pressure 

drop was assumed in calculating the mass flux and convective film coefficients resulting 

from circumferential flow in the joint. Figure 2.3.42 presents the predicted erosion and 

heat-affected rates in the joint. The erosion is insufficient to cause failure by losing the 

compressed portions of the nozzle-to-case insulation interface. 

The above analyses verify the thermal and structural integrity of the nozzleto-case insulation joint. Stress relief is shown to minimize both stresses and deflections 

in the joint. The stress relief component has a sufficient length to properly function 

throughout motor operation. In addition, the insulation joint is shown to be capable of 

withstanding the flow should it become vented and capable of maintaining a primary 

barrier that prevents gases from reaching the compressed Inconel C-ring and O-ring 

seals. 

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2.3.5 CASE SEGMENT AND CASE/NOZZLE JOINT SEALS 

The results of ARC studies of the subject seals are presented in this section. 

Highly reliable, commercially available, redundant, metal O-ring seals that meet all 

design requirements and are well suited to the ARC joint designs are recommended. 

2.3.5.1 JOINT/SEAL TRADE STUDY RESULTS 

As indicated in ARC'S Mid-Term Report (Appendix A), the primary goal of the 

joint/seal design effort was to improve the reliability of the system and minimize the 

possibility of failure. Based upon the excellent design and structural/thermal analysis 

results of NASA LaRC, ARC selected the concept of a bolted, flat flanged joint using 

face seals to meet this objective. Numerous seal designs (configuration and materials) 

were investigated as part of the case field joint trade study. The identification of the 

advantages of redundant metal seals was an important output of this effort. 

2.3.5.2 SEAL PRELIMINARY DESIGN CRITERIA 

The results of the trade study were utilized to establish the seal design criteria of Table 2.3.3. The listed parameters are rather straightforward with the exception 

of the following environmental values: 

Maximum Pressure - MEOP for the topmost case field joint is 1,004 psia; 

design pressure is 1.4 x 1,004 = 1,406 psia. The corresponding design 

pressure for the case/nozzle joint is 1.4 x 909 = 1,273 psia. For preliminary discussions with seal manufacturers, 1,4 10 psia design pressure 

was used for all case seals. 

Maximum Temperature - a value of 1,400"F was used. No physical significance was attached to the value; it enabled discussions with seal 

manufacturers without overly constraining them. 

Maximum Cap - a maximum gap (Le., flange face lift-off) of 3.5 mil was 

used. This value proved to be representative of actual results (a range of 

0 to 5.7 mils, as indicated in Section 2.3.5.4). 

Maximum Cap Rate - no a priori values of this parameter were available. 

(See Section 2.3.5.4.) 

It is to be noted that the seal criteria of Table 2.3.3 are preliminary in the 

sense that these are judged necessary but not sufficient conditions. In addition to the 

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TABLE 2.3.3. SEAL DESIGN CRITERIA 

CASE FIELD JOINTS AND 

NOZZLE/CASE JOINT. 

REDUNDANCY: TWO SEALS MINIMUM, VERIFIABLE BY TEST 

PRESSURE ASSISTANCE: SEALS TO ACCOMMODATE, BUT TO RETAIN SEALING 

CAPABILITY WITHOUT, PRESSURE ASSISTANCE 

DESIGN COMPATIBILITY: SEALS TO BE COMPATIBLE WITH BOLTED FLAT SEALING JOINT 

DESIGN AND ATTENDANT PLACEMENT AND TRACKING 

VELOCITY REQUIREMENTS PIS 

RELIABILITY: 

COMMERCIAL: 

ASSEMBLY: 

SEAL BEHAVIOR TO BE WELL UNDERSTOOD AND DEMONSTRATED 

OVER THE RANGE OF DESIGN OPERATION CONDITIONS 

DEMONSTRATED FABRICATION AND UTILIZATION HISTORY OF 

SIMILAR EQUIPMENT 

EASE OF INSTALLATION AND SEGMENT ASSEMBLY; RESISTANT 

TO DAMAGE; COMPATABILITY WITH BOLT TENSION 

REQUIREMENTS 

ENVIRONMENTAL: MAXIMUM PRESSURE (PSIA): 1410 

MAXIMUM TEMPERATURE (OF): 1400 

MINIMUM TEMPERATURE ( OF) : 20 

MAXIMUM GAP (MILS): 3.5(2) 

MAXIMUM GAP RATE: TBD(3) 

(l) SEALING CAPABILITY TO BE DEMONSTRATED AT TWO TIMES MAXIMUM GAP AT MAXIMUM 

TRACKING VELOCITY AND TWO TIMES TRACKING RATE AT MAXIMUM GAP. 

(2) A MAXIMUM OPENING OF THE JOINT SEALING SURFACES OF 3.5 MIL WAS 

ESTABLISHED BASED UPON PRELIMINARY WORST CASE STRESS ANALYSIS RESULTS. 

SEE SECTION 2.3.5.4 FOR FINAL RESULTS. 

(3) GAP RATES, WHICH ESTABLISH REQUIRED TRACKING RATE TO ENSURE SEALING, ARE 

PROVIDED IN SECTION 2.3.5.4. 

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parameters listed, numerous additional requirements exist relating to the manufacturing, 

transportation, assembly, and testing operations; joint design and performance (both 

static and dynamic) specifics provide additional seal requirements. Some of these additional requirements were derived from the analysis of joint performance and information 

from the seal manufacturers and are provided in Section 2.3.5.4. 

2.3.5.3 RELATED SEAL EXPERIENCE 

The use of metal seals in rocket motor applications has undergone an 

interesting evolution. A number of SRMs of the 1950 to 1960 timeframe incorporated 

"Flexatallic" type gasket seals. These seals are widely used, even today, in industrial and 

utility applications characterized by heavy flanges, high clamping (static) loads, and low 

dynamic loadings. The development of lighter, more highly stressed cases ended the use 

of these seals in rocket motor applications. Elastomeric O-ring seals were employed in 

the overwhelming majority of rocket motor case seal applications. Concurrently, aerospace manufacturers and their seal vendors were developing high performance metal 

seals. It is to be noted that Rocketdyne is in this group, and has made extensive use of 

advanced, pressure-assisted metal seals throughout the SSME. These seals have been 

found to provide very high reliability in high pressure service over a range from cryogenic liquid to high temperature gas applications. 

The widespread use of rubber O-rings in rocket motor applications is well 

documented. The catastrophic failure of the 51-L seals is, in and of itself, not sufficient 

to eliminate their use (especially when used with improved joint designs). Both metal and 

elastomeric seals were investigated in the trade study (Appendix A). Metal seals were 

found to have superior static and dynamic performance characteristics in this study. 

Metal seal superiority over rubber seals was notable in the area of reliability: breaching 

of the insulation seal would result in rapid and catastrophic loss of a rubber seal, while 

metal alloy seals offer much higher sealing reliability; low ambient (and resultant cold 

joint) temperatures do not result in metal seal impairment, whereas the static and 

dynamic sealing characteristics of rubber seals suffer severe impairment; metal seals 

offer improved analysis tractability/predictability over rubber seals. 

As indicated in the trade study (Appendix A), the metal-jacketed gasket 

primary and metal C-ring secondary seal configuration was highly rated for the joint 

designs investigated. The ARC modified in-line joint with this seal configuration was 

reviewed by the seal manufacturer with extensive design and manufacturing background 

in providing seals to similar requirements. In summary, the Fluorocarbon Gasket 

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Division* did not recommend (and declines to bid) the use of a gasket type seal and both 

the Fluorocarbon Components Division and the Advanced Products, Inc. recommended 

the use of silver plated, Inconel 718 O-rings for both primary and secondary seals. These 

two manufacturers share the supply of seals for utility nuclear reactor vessel head 

closures. Their manufacturing considerations and recommendations coupled with relevant utilization experience resulted in prime consideration being given to a dual metal 

O-ring seal at case segment joint and case-nozzle joint locations. 

Design data for various metal seal designs were provided by Advanced Products, Inc. and are included in Tables 2.3.4 through 2.3.7. Table 2.3.4 provides yield 

strength versus temperature data for a number of seal alloys; the superior high temperature strength characteristics of age-hardened Inconel 7 18 is apparent. Design pressure 

ranges for various seal designs are included as Table 2.3.5; as indicated, vented metal 

O-ring nominal design range is from loe6 Torr. to 3,000 psia (standard, i.e., 10 percent, 

wall thickness). Nominal design temperature upper limit for Inconel 718 is 1,400'F as 

indicated in Table 2.3.6. Recommended materials for various seal configurations are also 

indicated in Table 2.3.6. Seal plating maximum working temperatures and recommended 

sealing loads for a variety of materials are provided in Table 2.3.7. A design temperature ceiling of 1,500"F and a minimum flange loading of 250 pounds/circumferential-inch 

are indicated for silver plating. 

2.3.5.4 SEAL DESIGN AND PERFORMANCE SPECIFICS 

Based upon a review of the ARC preliminary joint design and performance 

requirements, Fluorocarbon, Inc. recommended the following seals: 

RANKING DESCRIPTION 

HIGHEST 

SECOND 

THIRD 

2 METAL** O-RINGS, 

2 METAL** O-RINGS, 

2 METAL** C-RINGS, 

I /P DIAMETER 

3/8" DIAMETER 

3/8" DIAMETER 

* Fluorocarbon, Inc., Metallic Gasket Division of Houston, Texas; Fluorocarbon, Inc., 

Components Division of Columbia, S.C.; and Advanced Products, Inc, of North 

Haven, Connecticut. 

** Inconel 718 with silver plating, 10 percent wall 

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TABLE 2.3.4. METAL SEAL MATERIAL DESIGN DATA 

(ADVANCED PRODUCTS, INC. ) 

Weld Strength (psi) for Various Materials at Various Temperatures 

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(ADVANCED PRODUCTS, INC. ) 

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DESIGN 

RANGE FOR 

METAL I 

ED ! 

f_ 

NG 

VEN 

0- R 

LEGEND: 

A = STANDARD WALL 

B = HEAVY WALL 

V = VENTED 

M-W-R = WIRE RING 

M-0-R = O-RING 

M-V-R = V-RING 

M-C-R = C-RING 

M-SE-R SPRING-ENERGIZED RING 

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M-W-R M9R M-V-R M-CR MSETABLE 2.3.6. METAL SEAL MAXIMUM TEMPERATURE DESIGN DATA 

(ADVANCED PRODUCTS, INC. ) 

brtelloy 

Cold 

Copper 

Inconel718 

Inconel-X-750 

InconcldOO 

Aluminum 

300Scrier-SS 

1 BOO’F 

1700°F 

1700’F 

14WF 0 

1400°F e e 

1200°F 

BOO’F 

4OO’F 

Nickel 

Hayncr 25 0 

Maximum 

Working 

Temperature 

1WF 

450” F 

300°F 

1500°F 

17W F 

1700” F 

2300°F 

Rscomnrmded Lord 

PMing’Corting Range tor Sealing 

Typc (Ibr.lclrwm. in.) 

Indium 75-350 

Teflon 150-450 

bad 100400 

SWl 250 up 

Copper 250 up 

Gdd 200 UP 

Nickel 400 up 

LEGEND : 

M-W-R = WIRE RING 

M-0-R = O-RING 

M-V-R = V-RING 

M-C-R = C-RING 

M-SE-R = SPRING-ENERGIZED RING 

TABLE 2.3.7. METAL SEAL LOADING AND PLATING DESIGN DATA 

(ADVANCED PRODUCTS, INC . ) 

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The springback characteristics for the 3/8-inch outside diameter O-ring are 

included as Figure 2.3.43. This figure presents load versus compression data for a 4-inchlong specimen. The characteristic indicated as "20% COMP" corresponds to a flange 

groove height of 0.300 inch, Le., 80 percent of a 0.375-inch diameter; the resultant 

compression is 0.075 inch. The springback for this configuration is 33 mils, i.e., the 

difference between the 75-mil compression and the 42-mil permanent set of the ring. 

The springback value of 12 mils quoted by Fluorocarbon, Inc., therefore, incorporates a 

good margin of conservatism. In addition, the 12-mil quoted springback is more than 

twice the worst-case nozzle-to-case joint gap of 5.7 mils and nearly four times the 

3.1-mil maximum gap at any case segment field joint. 

The 3/S-inch-diameter O-ring configuration was selected by ARC since it 

requires less room and therefore enables a smaller groove (and hence flange); the smaller 

O-ring also required a smaller compression load (about 80 percent of the l/Z-inch ring); 

at 0.012 inch, the springback for the 3/8-inch rings is, in all cases, more than twice the 

maximum design flange gap. While requiring higher flange loading and offering less 

springback than the comparable C-ring designs, the O-ring offers more seal area and is 

more forgiving to distortion in seal or flange bearing surface. 

The revised Contract End Item (CEO requirement is that "each seal shall 

demonstrate sealing capability at two times maximum gap size at maximum expected 

tracking velocity and two times tracking rate at maximum expected gap site." Neither 

of the seal manufacturers were able to provide tracking r a t e d a t a or mathematical model 

results for any metal seals. The extremely rapid response rate of the metal O-ring seals 

had negated the manufacturers' need for such information; both Fluorocarbon and 

Advanced Products did, however, indicate a willingness to discuss test programs to 

acquire the information. 

Lacking this tracking rate data, two mathematical modeling efforts were 

initiated. A highly simplified model establishing a "seal characteristic" based upon 

calculated natural frequencies and resultant average cyclic velocities was prepared. The 

resultant envelope was then "halved" to establish an "operating ceiling" as per the above 

CEI requirement. These results were then plotted along with the joint operating maps 

from the detailed structural analysis efforts. These results are included in Figures 2.3.44 

and 2.3.45 for the case/nozzle and case field joints, respectively. As indicated, due to 

rhe very high seal tracking rates and large springback, seals at both locations provide 

wide "operating margin" even at worst-case conditions. 

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FIGURE 2.3.44. CASE/NOZZLE JOINT AND SEAL CHARACTERISTICS. 

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FIGURE 2.3.45. CASE FIELD JOINT AND SEAL CHARACTERISTICS. 

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The joint structural model dynamic results provided the joint characteristics 

presented in Figures 2.3.44 and 2.3.45. Detailed results are presented in Sections 2.3.1 

and 2.3.3 for the case field joint and case/nozzle joints, respectively. 

The design and performance specifics for the dual metal O-ring seals are 

provided in Table 2.3.8. Backup information on the seals is included in Appendix B. 

2.3.5.5 SEAL SUMUARY 

A review of the dual metal O-ring seal design and performance information of 

the previous section with the criteria of Table 2.3.3 reveals that the design conforms 

with requirements in each area. In the evaluation of this design, the demonstrated 

fabrication and utilization history for seals of this type is a very important consideration. Also most important is the fact that these seals lend themselves well to the joint 

designs. Finally, the selection of the design followed from the review and recommendations of the leading seal manufacturers. 

Among candidate topics for further study are design, fabrication, shipping, 

assembly, inspection, etc. Careful review by ARC and the leading suppliers of these 

seals has led ARC to the conclusion that, for the system proposed, each of these 

obstacles can be overcome in straightforward, cost-effective fashion with a high degree 

of reliability. 

2-87 
TABLE 2.3.8. SEAL DESIGN DATA. 

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JOINT 

CASE SEGMENT 

CASE SEGMENT 

CASE/NOZZLE 

CASE/NOZZLE 

NOTES 

- SEAL 

PRIMARY 

SECONDARY 

PRIMARY 

SECONDARY 

NOMINAL 

SEAL OD (IN) 

141.1 

142.4 

103.1 

104.4 

GROOVE OD EQUALS SEAL OD + 0.014/.029. 

GROOVE HEIGHT AND WIDTH ARE 0.300 INCH AND 0.445 INCH 

RESPECTIVELY AT ALL LOCATIONS. 

ALL SEALS ARE 3/8-INCH DIAMETER, INCONEL 718, SILVER PLATED 

(0.004/0.006 INCH), 0.038-INCH WALL, VENTED DESIGN. 

REQUIRED SEAL COMPRESSION IS 2500 LBF/IN-SEAL (FLUOROCARBON 

VALUE VS. 1700 ADVANCED PRODUCTS). 

DESIGN SPRINGBACK IS 0.012-INCH (FLUOROCARBON VALUE VS. 0.011 

ADVANCED PRODUCTS). 

SEAL SURFACES TO BE 32 MICRO-INCHES (RMS) FINISH; MACHINE TOOL 

MARKS TO BE CONCENTRIC. LUBRICATION NOT REQUIRED. 

SEALS NOT TO BE REUSED. 

2-88 
-~ ~ ~~ 

2.4 ASBESTOS-FREE INSULATION 

The insulation trade studies were presented in the Mid-Term Report, which is 

submitted as Appendix A of this document. These trade studies resulted in an asbestosfree insulation design for the Block I1 SRM. This section will present this overall insulation design and the supporting rationale. In addition, the available material database will 

be presented followed by the thermal analyses that have been performed since the 

Mid-Term Report was issued. These analyses were used to verify the adequacy of the 

materials selected to achieve the desired insulation performance safety factors. 

The specific design descriptions and rationale for the insulation at the case 

segment joints and the case-to-nozzle joint were presented in Sections 2.3.2 and 2.3.4, 

respectively. 

2.4. I INSULATION DESIGN DESCRIPTION 

The case insulation trade studies resulted in selection of a hybrid insulation 

system for the case segments. This system uses a Kevlar/silica-filled Hypalon as the 

material used next to the steel case wall to provide the required thermal protection. The 

aft portion of the motor case segments near the joints and the portion of the aft case 

segment that is exposed during much of the motor firing have a layer of USR-3800 NBR/ 

phenolic insulation placed between the Hypalon and the propellant combustion gas path. 

This material provides high erosion resistance in these critical areas in the aft case and 

at the joints. The Mid-Term Report described the automated ribbon winding technique 

that has been chosen as the best method for applying the insulation into the case segments. 

USR-3800 has also been selected as the material for the molded inhibitors 

located at the forward end of each casting segment. These inhibitors would be formed 

during the case segment insulation automated layup. 

The propellant liner and the castable inhibitors on the aft end of the grain 

segments will be made from the same material. The trade studies performed during the 

first phase of the work identified that the best option would be to retain the current liner 

material and replace only the asbestos filler. This choice would keep intact the proven 

compatibility of the CTPB liner with the PBAN propellant and provide known bonding and 

2-89 
2-90 

aging characteristics. A number of fibrous materials, including silicates similar to 

asbestos, were identified as candidates. These include PBI, carbon fibers, ceramic fibers, 

silica fibers, and mineral wool. The advantages and disadvantages of these fillers and a 

relative ranking are shown in Table 2.4.1. 

Carbon fiber was given the highest score due to its ability to maintain maximum char strength at high temperatures, although to achieve maximum strength it must 

be used in conjunction with a silicate that can melt and inhibit surface oxidation. All the 

fibrous silicate materials behave similarly in the high temperature environment and 

cannot be differentiated on that basis. A laboratory study would be needed to differentiate on the basis of processing and liner mechanical properties. PBI fiber was given the 

lowest score due to lack of an experience base. 

The selected liner retains the current binder system with the asbestos being 

replaced 50 percent by carbon fiber and 50 percent by a silicate fiber. Cab-0-Si1 

(another silicate) would be added to control the thixotropic behavior for processing 

purposes. 

As discussed in the Mid-Term Report, replacement materials for the 

asbestos-containing portions of the safe and arm (S&A) device were also identified. The 

S&A clutch disc material will be replaced with a Kevlar phenolic, while the S&A commutator material will be replaced with ceramic phenolic. 

A silica NBR material will be used to fabricate the ssress relief flaps at the 

aft end of each casting segment and to form a component of the joint insulation. This 

material was selected because of the need for a low modulus and compatibility with the 

PBAN propellant. 

2.4.2 INSULATION MATERIAL DATABASE 

USR-3800 is an existing NBRlphenolic material that uses boric acid as a 

filler. USR-3800, a Uniroyal product, has been produced for close to 20 years and was 

used in both the POLARIS and POSEIDON missiles. In addition, the material has been 

tested as part of insulation studies for large motor applications (References 1 and 2) and 

been found to exceed the performance of the currently used NBR/asbestos material. 

Table 2.4.2 presents the thermal and mechanical properties of this material that have 
TABLE 2.4.1. LINER MATERIAL ALTERNATIVES. 

MATERIAL/MANUFACTURER 

CARBON FIBER 

BASF, HITCO, HERCULES, 

UNION CARBIDE, CORTAULDS 

CERAMIC FIBER (SILICATE) 

NEXTEL - 3M 

FIBERFRAX - H.I. THOMPSON 

FIBROUS SILICA 

REFRASIL - H.I. THOMPSON 

MINERAL WOOL (SILICATE) 

PBI FIBER 

CE LANE SE 

ADVANTAGES 

DEMONSTRATED IN INSULATION 

fORMULATIONS TO MAINTAIN 

EXCELLENT CHAR STRENGTH; 

DOESN'T MELT 

MELTS AND PROVIDES CORROSION PROTECTION 

MELTS AND PROVIDES CORROSION PROTECTION 

MELTS AND PROVIDES CORROSION PROTECTION 

DEMONSTRATED IN INSULATION FORMULATIONS TO GIVE 

GOOD HIGH TEMPERATURE 

PERFORMANCE. 

2-91 

DISADVANTAGES - SCORE 

CAN BE OXIDIZED AT 7.5 

HIGH TEMPERATURES 

HIGHER COST 7.0 

7.0 

7.0 

NO PROCESSING DATA 5.0 

AVAILABLE ON LINER 

FORMULATIONS. LONG 

ISTICS NOT ESTABLISHED. 

TERM AGING CHARACTER- 
TABLE 2.4.2. USR-3800 MECHANICAL AND THERMAL PROPERTIES. 

TEMPERATURE SPECIFIC HEAT CONDUCTIVITY 

("K) ("R) (kJ/kq-K) JBtulLbm-R) (W/rn-K) (Btu/ft-sec-Rl 

256 460 0.711 0.170 0.031 0.50 E-5 

28 3 510 1.091 0.261 0.083 1.33 E-5 

311 5 60 1.463 0.350 0.138 2.22 E-5 

339 610 1.831 0.438 0.167 2.68 E-5 

367 660 2.044 0.489 0.175 2.81 E-5 

39 4 710 2.136 0.511 0.180 2.89 E-5 

422 760 2.153 0.515 0.182 2.92 E-5 

MECHANICAL PROPERTIES AT 

AMBIENT CONDITIONS 

STRESS STRAI N MODULUS 

JMPal (PSI) 0 lMPa1 JPSI) 

7.6 1100 200 179.3 26,000 

2-92 
been obtained from the literature. Figure 2.4.1 presents the results of a thermogravimetric (TGA) analysis of both USR-3800 and the current NBR/asbestos material. The 

TGA analysis is used to determine the decomposition characteristics of the material. As 

can be seen from the figure, the residual char material for USR-3800 is greater than that 

for NBR/asbestos, which is an indication of improved char strength. Table 2.4.3 shows 

data from Reference 1, which compare the erosion characteristics of USR-3800 and 

NBR/asbestos. As can be seen from the data, USR-3800 outperforms NBR/asbestos. 

These data were used to create an analytical model of this material to aid the thermal 

analysis discussed in Section 2.4.3. All additional required data needed to validate the 

use of this material will be obtained in the Development and Verification Phase of the 

Block I1 program. 

The Kevlar/silica-filled Hypalon material is a recently developed material 

currently being used by ARC as an asbestos replacement insulation in a surface-to-air 

missile system. This rocket motor is currently in the qualification phase, Performance, 

bond, and aging data have been and are being generated in this program and will be 

available for the Block I1 Program. Table 2.4.4 presents the thermal and mechanical 

properties of this material. Figure 2.4.2 presents the results of a TGA analysis of the 

Kevlar/silica-filled Hypalon. A validated thermal model of this material existed before 

this current study program began and has been used to verify the adequacy of the proposed design. 

Table 2.4.5 contains the mechanical and thermal properties of the silica NBR 

material selected for the stress relief flaps and a joint insulation component. 

The trade studies performed indicated that a hybrid insulation system using a 

Kevlar-filled EPDM material in place of the USR-3800 also was a viable alternative to 

the current NBR/asbestos insulation. ARC believes that it is good design practice to 

carry a backup insulation material into the Development and Verification Phase of the 

program. Kevlar-filled EPDM is recommended as a backup insulation material for the 

USR-3800. ARC has an established database with the Kevlar-filled EPDM material 

(References 3 and 4). Table 2.4.6 presents the thermal and mechanical properties of this 

material. Figure 2.4.3 shows the results of a TGA analysis of Kevlar-filled EPDM. 

Figure 2.4.4 shows the type of erosion data that have been obtained using the ARC Insulation Test Motor (ITM). As can be seen from the figure, both instantaneous and average 

erosion data as a function of Mach number are available for use in the design process. 

2-93 
FIGURE 2.4.1. THERMOGRAVIMETRIC ANALYSIS RESULTS FOR 

USR-3800 AND NBR/ASBESTOS. 

2-94 
TABLE 2.4.3. COMPARISON Of EROSION DATA FOR USR-3800 AND 

NBR/ASBESTOS . 

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EROSION RATE DATA^ 

INITIAL USR-3800 NBR/ASBESTOS 

MACH NUMBER (mm/sec) ( i n/sec) (mm/sec) (i n/sec) 

0.05 

0.10 

0.15 

0.20 

0.25 

0.102 0.004 0.229 0.009 

0.152 0.006 0.508 0.020 

0.076 0.003 0.660 0.026 

0.076 0.003 0.787 0.031 

0.051 0.002 0.864 0.034 

DATA FROM REFERENCE 2.4.1. PROPELLANT HAD 15% ALUMINUM, 88% SOLIDS. 

AVERAGED EROSION BASED ON INITIAL MACH NUMBER. 

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TEMPERATURE 

("K)("R) 

200 360 

294 530 

422 760 

644 1160 

811 1460 

1367 2460 

2478 4460 

3589 6460 

TABLE 2.4.4. HYPALON MECHANICAL AND THERMAL PROPERTIES. 

SPECIFIC HEAT 

(kJ/kq-K) (Btu/Lbm-R) 

0.732 

1.484 

1.848 

1.986 

2.048 

2.090 

2.090 

2.090 

0.175 

0.355 

0.442 

0.475 

0.490 

0.500 

0.500 

0.500 

CONDUCT IV ITY 

(W/m-K] (Btu/ft-sec-R) 

0.122 

0.112 

0.099 

0.077 

0.069 

0.069 

0.069 

0.069 

1.95 E-5 

1.80 E-5 

1.59 E-5 

1.23 E-5 

1.10 E-5 

1.10 E-5 

1.10 E-5 

1.10 E-5 

MECHANICAL PROPERTIES AT 

AMBIENT CONDITIONS 

STRESS STRAIN MODULUS 

(MPa) (PSI) 0 (MPa) (PSI) 

11.0 1600 28 58.6 8,500 

2-96 

EMISSIVITY 

0.80 

0.80 

0.80 

0.80 

0.80 

0.80 

0.80 

0.80 
ORIGINAL PAGE IS 

OF POOR QUALITY 

F IGURE 2.4.2. THERMOGRAV IMETR IC ANALYSIS RESULTS FOR 

KEVLAR/SIL ICA-FILLED HYPALON. 

2-97 
TABLE 2.4.5. SILICA NBR MECHANICAL AND THERMAL PROPERTIES. 

TEMPERATURE SPECIFIC HEAT CONDUCT1 VITY 

("K) ("R) JkJ/kg-K) (Btu/Lbm-R) (W/m-K) (Btu/ft-sec-R) 

256 460 1.818 0.435 0.242 3.89 E-5 

283 510 1.881 0.450 0.242 3.89 E-5 

311 5 60 1.923 0.460 0.242 3.89 E-5 

339 610 1.986 0.475 0.248 3.98 E-5 

367 660 2.048 0.490 0.260 4.17 E-5 

394 7 10 2.111 0.505 0.262 4.21 E-5 

422 7 60 2.174 0.520 0.267 4.28 E-5 

MECHANICAL PROPERTIES AT 

AMBIENT CONDITIONS 

STRESS STRA I N MODULUS 

(MPa) (PSI) 0 JMPa) IPS I ) 

13.5 1960 470 14.8 2150 

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~' TABLE 2.4.6. KEVLAR/EPDM MECHANICAL AND THERMAL PROPERTIES. 

TEMPERATURE SPECIFIC HEAT CONDUCT I V I TY 

("K) ("R) JkJ/kg-K) (Btu/Lbm-R) (W/rn-K) (Btu/ft-sec-Rl EMISSIVITY 

217 390 

256 460 

350 630 

422 760 

644 1160 

722 1300 

1367 2460 

3589 6460 

1.038 

1.170 

1.492 

1.714 

1.756 

1.756 

1.756 

1.756 

0.2483 

0.2800 

0.3570 

0.4100 

0.4200 

0.4200 

0.4200 

0.4200 

0.196 

0.190 

0.162 

0.162 

0.125 

0.112 

0.112 

0.112 

3.15 E-5 

3.05 E-5 

2.60 E-5 

2.60 E-5 

2.01 E-5 

1.80 E-5 

1.80 E-5 

1.80 E-5 

MECHANICAL PROPERTIES AT 

AMBIENT CONDITIONS 

STRESS STRAIN MODULUS 

(MPa) (PS 11 0 JMPa) IPS11 

7.8 1125 19 72.4 10,500 

2-99 

.90 

.90 

.90 

.90 

.90 

.90 

.90 

.90 
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FIGURE 2.4.3. THERMOGRAVIMETRIC RESULTS FOR KEVLAR FILLED EPDM. 

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Gd : R I I w I I t t’ 

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WCH NCM6GIR 

FIGURE 2.4.4. EROSION DATA FOR KEVLAR-FILLED EPDM OBTAINED 

FROM THE ARC INSULATION TEST MOTOR. 

2-101 

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2.4.3 CASE INSULATION THERMAL ANALYSES 

One-dimensional thermal analyses were conducted on the center segment of 

the SRM motor. This section was chosen because performance data were available at 

this location from the Morton Thiokol Inc. thermal report on the high performance motor 

(HPM) design (Reference 5). The Aerotherm Chemical Equilibrium (ACE) and the 

Charring Material Thermal Response and Ablation (CMA3) computer programs were used 

in these analyses. These programs are standard industry tools for insulation analysis. 

ACE is used to perform a series of thermochemical calculations involving 

varying mixtures of propellant combustion gas, insulator pyrolysis gas, and insulator char 

constituents. The thermochemical calculations quantitatively describe the equilibrium 

reaction state (i.e., species, temperature), chemical energies, and energies associated 

with mass transfer occurring at the heated surface of the insulator for each mixture. 

CMA3 uses this matrix of energy descriptions in performing a heated surface 

energy balance. CMA3 incorporates this surface energy balance in a finite difference 

solution of the one-dimensional, axisymmetric parabolic heat conduction equation which 

calculates transient in-depth temperatures, erosion recession rates, pyrolysis gas generation rates, and the density gradient within the insulator. Remaining parameters required 

by CMA3 include the geometric configuration, material properties, and boundary conditions. The internal boundary conditions consist of three components: convection, chemical energy, and luminous radiation. The convective conditions were identical to those in 

Reference 5. 

The thermal analyses performed on the center motor segment used the current HPM insulation thicknesses so that a direct comparison of the resulting safety 

factors could be made to the current asbestos-containing material. The portion of the 

overall thickness allotted to the Hypalon material was that required for thermal protection of the case. The remainder was analyzed as USR-3800. The results of these 

analyses are presented in Figure 2.4.5. As can be seen from the figure, the USR-3800/ 

Hypalon hybrid design outperforms the current NBR/asbestos material, resulting in 

higher safety factors for the same insulation thickness. 

Because of differences in the specific gravity of the selected insulation 

materials when compared to the current NBR/asbestos, a weight savings of 806 kg (1,774 

2-102 
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CURRENT CENTER SEGMENT HPM DESIGN (REF. 2.4.5) 

I I 1 1 I I 

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0.0 8.0 16.0 24.0 32.0 40.0 48.0 0’ 

DISTANCE FROM AFT END OF SEGMENT (IN.) 

F IGURE 2.4.5. THERMAL ANALYSIS SAFETY FACTOR COMPARISON 

FOR THE CENTER SEGMENT. 

2-103 
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pounds) over the HPM design has been calculated when insulation thicknesses are held at 

the current values. 

Because of the improved safety factors for the proposed design, the potential 

exists for the Block I1 design to thin the insulation thickness and obtain a larger volume 

for propellant loading while maintaining the current insulation safety factors. This 

reduction in insulation thickness reduces the insulation weight an additional 1,163 kg 

(2,559 pounds). The ballistic performance impact of this weight reduction is discussed in 

Section 3.0. 

In summary, the insulation materials selected as a result of the trade studies 

have been found to have an adequate database of properties. These properties have 

permitted the creation of thermal analytical models that have been used to analyze the 

proposed Block I1 design. These preliminary analyses have established that the selected 

nonasbestos materials are viable replacements for the current asbestos materials. The 

additional data necessary to finalize the Block I1 design will be obtained in the Development and Verification Phase of the program. 

2-104 
~~ 

REFERENCES 

1. 

2. 

3. 

4. 

5. 

Simmons, B. A., and D. L. Nachbar, "Development of Cost-Optimized Insulation System for Use in Large Rocket Motors," Vols. 1-111, Aerojet-General 

Corporation, NASA CR-72581, CR-72582, and CR-72583, 1969. 

Bradley, W., J. DeAcetis, A. Stenersen, et al., "Investigation and Evaluation 

of Motor Insulation for Multiple Restart Application," Final Report, AFRPLTR-67-287, 1967. 

Yezzi, C. A,, and B. 8. Moore, "Characterization of Kevlar/EPDM Rubbers 

for Use as Rocket Motor Case Insulators," AIAA Paper 86-1489, 1986. 

Spear, G. B., T. L. Ludlow, J. F. Sparks, B. B. Moore, and T. F. Davidson, 

Thermally Stable Elastomeric Insulations for Tactical Applications," presented at the JANNAF Propulsion conference, August 1986. 

Eddy, N. F., K. Eckhart, S. B. Pendleton, et al., "High Performance SRM 

Internal Insulation Design Report," Morton Thiokol, Inc., TW R-13065 Rev. B 

Amendment 2, 1984. 

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2.5 NOZZLE DESIGN 

At the beginning of the Block I1 SRM Study Program, the decision was made 

to retain the current nozzle concept to take full advantage of the years of design, development, and testing invested in that approach. Thus, alternative approaches to the basic 

nozzle geometry and assembly techniques, the flex bearing, and the thrust vector control 

system were not considered directly, Impacts in these areas were considered only when 

driven by the results of the study in other areas. One item identified for study was the 

nozzle liner material, which has exhibited some anomalous erosion in the past (STS-8). In 

addition, the Block I1 requirements dictated that redundant seals be included at nozzle 

internal assembly joints and that asbestos-containing materials be eliminated. The 

results in each of these study areas is discussed in the following section. 

2.5.1 NOZZLE LINER MATERIAL 

Liners used on the current SRM nozzle are fabricated from rayon-based 

carbon cloth/phenolic tape. The performance of this liner material has been adequate 

with the exception of anomalous gouging and "pocketing erosion", which was first noted 

on STS-8. NASA and the current nozzle fabricator have spent considerable effort investigating the solutions to this problem and, as a result of stringent material/processing 

controls and revised processing/fabrication techniques, have virtually eliminated the 

incidence of pocketing erosion in the six flights preceding 51-L. 

The goal of the program study in this area was to investigate the use of stateof-the-art liner materials to further reduce the possibility of pocketing erosion. A 

material trade study was, therefore, initiated and the preliminary results were given in 

the Mid-Term Report (Appendix A). Since that time, the study was completed and the 

final results are included as an addendum to Appendix A. Contrary to the preliminary 

findings, the final conclusion was that the current material should be retained. In short, 

none of the materials investigated meet the requirements of offering a positive gain in 

reliability and having an extensive database to back them. The most promising material 

was found to be multidirectionally reinforced carbon-carbon advanced composites. While 

carbon-carbon composites are most certainly the material of the future in this 

application, it was not yet considered sufficiently mature for the shuttle SRM. Refer to 

the addendum to Appendix A for the details of the trade study and, in particular, for the 

results of a design study submitted by Societe Europeenne de Propulsion (SEP), a leading 

2-106 
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French propulsion company. SEP investigated two carbon-carbon ITE approaches utilizing nondegradable backup insulative materials, which are certainly viable options for the 

future, particularly in applications where performance requirements dictate the use of 

the most advanced materials available. 

2.5.2 NOZZLE SUBASSEMBLY JOINT SEALS 

A redundant seal design study was conducted for the five joints in the nozzle 

subassembly shown in Figure 2.5.1. The details of that study are included in Appendix 

A. As reported there, the selected joint/seal configuration is the face seal and bore seal 

combination. The selected design at each joint is shown in Figure 2.5.2. 

The seals chosen for use at each location are elastomeric O-rings. This type 

of seal is currently used; it appears to have performed adequately in this application in 

the past, as no record of any problems has been found. In addition, potential joint opening that dictated the design at the case and nozzle-to-case joints is less of a problem for 

these joints, which are typically in compression and are not subject to the load eccentricities of the two "major" joints. As previously reported, the incorporation of redundant scale at nozzle subassembly joints result in 247.5 pounds. 

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2.5.3 FLEX BEARING BOOT 

Replacement of the silica and asbestos-filled NBR in the nozzle flex bearing 

boot presents numerous design challenges due to the unique structural and thermal requirements imposed upon it. Low flexural stiffness is mandatory to minimize thrust 

vector control actuation torque. The current stacked wafer approach using carbon cloth 

between wafers (for venting purposes) helps achieve the low flexural stiffness. Replacement materials must be able to be processed in the same manner. Boot deflections 

resulting from pressurization and thrust vectoring stress the char and may cause char 

spallation, which increases the affected rates. 

Thermal environments imposed on the boot are severe due to the configuration and location. The bent or recessed portion of the boot can trap aluminum oxide 

particles in the combustion gas. Heat is rapidly conducted from the aluminum oxide to 

the boot. In addition, flow of aluminum oxide along the wall will cause gouging of the 

boot. 

Candidate materials with a sufficient database for the boot include NBR 

filled with silica; silica and polybenzimidazole (PBI) fiber; or silica and Kevlar fiber. 

Other candidate materials are USR-3800 (NBR/phenolic), Kevlar-filled EPDM, and 

DC 93-104. Basic properties of these elastomeric insulations are presented in Table 

2.5.1. 

Each of the parameters influencing the operation of the flex bearing boot 

must be properly accounted for to select the optimum material. Thus, a trade study was 

performed to evaluate the relative performance of candidate materials. Relative 

weighting factors used in the trade study emphasized reliability as the most important 

parameter. Table 2.5.2 shows the results of the trade study. DC 93-104 is shown to have 

properties that best satisfy the requirements for the flex bearing boot. 

As previously mentioned, DC 93-104 has a strong char that adheres well to 

the virgin insulation. This strong char is fairly brittle and can spa11 when subjected to 

significant deflections. To alleviate spallation problems, a through-the-thickness braided 

structure impregnated with DC 93-104 is used in the first three outer layers of the 

boot. Carbon fibers in the braided structure provide an integral retention mechanism for 

the char layer and have sufficient play to accommodate boot deflections. Figure 2.5.3 

shows the basic design details of the flex bearing boot. 

2-110 
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2-113 
2.5.4 ADHESIVES 

Two asbestos-filled adhesives used in the nozzle assembly are EA-934 and 

EA-91 3. Adhesive replacement studies performed on various programs at ARC resulted 

in selecting nonasbestos versions of the above adhesives, designated as EA-934 NA and 

EA-913 NA. Table 2.5.3 presents a comparison of basic properties between the current 

and the asbestos-free adhesives. 

TABLE 2.5.3. ADHESIVE PROPERTIES. 

EA-934 EA-934 NA EA-9 13 EA-913 NA 

VISCOSITY (POISE) AT 24°C 

PART A 7500 1760 3000-6500 2000-8000 

PART B 30 30 3-6 3-6 

TENSILE SHEAR STRENGTH* (MPA) 

24°C 21.3 24.6 30 34 

149°C 6.9 10.0 - - 

204°C 5.2 4.6 - - 

POTLIFE (MINUTES) AT 24°C 40 40 480 480 

* BONDED TO CHROMIC ACID ETCHED 2024-T3 ALCLAD. 

2-114 
2.6 IGNITER 

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2.6.1 INTRODUCTION 

The primary purpose of the igniter design study was to minimize potential 

leak paths in the igniter-to-adapter and adapter-to-motor case interfaces. Secondary 

objectives were to evaluate igniter propellant and ballistic design, evaluate expendable 

versus reusable hardware, and replace the igniter insulation with an asbestos-free 

material. 

The constraints on this design study were to maintain ignition performance 

and reproducibility without degrading structural and thermal margins. In order to have a 

high degree of confidence in any design changes, the database for the design change must 

exist or be created through a test program. 

2.6.2 CURRENT SRM IGNITER DESIGN 

The current shuttle igniter design is depicted in Figure 2.6.1. The SRM 

ignition system is a forward end, internally mounted solid rocket type (pyrogen) igniter 

and is approximately 44.5 inches long by 20 inches in diameter. The flight grain is a 

40-point star configuration approximately 16.9 inches in diameter by 32.8 inches long. 

The propellant grain consists of approximately 137 pounds of a 10 percent aluminized 

PBAN propellant and it is cast into a D6AC steel case insulated internally and externally 

with asbestos and silica-filled NBR. A molded silica phenolic throat insert controls the 

igniter pressure and directs the igniter plume to the main SRM propellant grain. 

The igniter chamber is bolted to the igniter adapter (D6AC steel) with 36 

3/4-inch bolts. Each bolt uses a special washer and pressure sealing packing. The main 

seal between the igniter chamber and the igniter adapter is a dual 0-seal gasket. The 

adapter bolts to the main SRM chamber with 40 5/8-inch bolts utilizing a washer and 

pressure sealing packing on each bolt. The primary seal between the adapter and the 

SRM chamber is also a dual 0-seal gasket. 

The ignition initiator is a small, multinozzled asbestos- and silica-filled NBR 

insulated steel cased rocket motor containing 1.4 pounds of propellant in a 30-point star 

configuration. The initiator case and the safety and arming (SdrA) device attach to the 

2-115 
t 

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Or POOR UJALITY 

FIGURE 2.6.1. CURRENT IGNITER AND SEALS. 

2-116 
igniter adapter. The S&A is bolted to the adapter using 10 bolts. A dual 0-seal gasket 

forms the dual redundant seals with special packing on each bolt as an environmental 

seal. 

The SdcA device consists of a reusable actuating and monitoring (AhM) and an 

expendable booster-barrier assembly containing a mixture of BKN03 pellets and granules. Two redundant NASA standard initiators (NSIs) provide positive ignition. The NSIs 

utilize dual redundant seals and the AdcM uses dual O-ring seals on the barrier rotor 

shaft. 

Totaling up the seals in the igniter/S&A, there are 6 primary seals, 42 

secondary seals, and 88 environmental seals. The primary seals are the fundamental 

seals that hold igniter or motor gas pressure while secondary seals would seal against gas 

pressure only if the primary seals failed. The environmental seals are used for sealing 

out the environments except for the bolt seals on the bolts that attach the igniter chamber to the igniter adapter. These seals are secondary seals and environmental seals. 

2.6.3 IMPROVED SRM IGNITER SYSTEM 

The improved SRM igniter is depicted in Figure 2.6.2. This system is a forward end, internally mounted solid rocket type (pyrogen) igniter. The igniter is approximately 19 inches in diameter by 34 inches long, overall. The flight grain is a 40-point 

star configuration that is 16.4 inches in diameter by 21.6 inches long. The propellant 

grain consists of 119 pounds of 18 percent aluminized HTPB propellant cast into a 200 

maraging steel case with an integral welded igniter adapter and a removable aft closure 

held in place using 36 high strength 3/4-inch bolts. The case is insulated internally and 

externally with Kevlar and silica-filled Hypalon. A molded silica phenolic throat insert 

controls the igniter pressure and directs the igniter plume to the main SRM propellant 

grain. 

The igniter adapter is bolted to the main SRM chamber with 40 5/8-inch bolts 

that have a washer and an environment seal on each bolt. The primary seal consists of a 

radially compressed aerospace C-T ring that seals against high pressures with larger 

clearances than an O-ring. This design is utilized for dynamic rod and piston seals and 

will not twist under installation. The secondary seal is a resilient metal C-ring mounted 

as a face seal between the adapter and the main SRM chamber. Metal C-ring seals are 

2-117 
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OF POOR QUALITY 

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2-118 
very high temperature seals (up to 2,000"F) and they are resilient seals that maintain 

sealing in the event of flange separation caused by thermal or pressure shock or by bolt 

creep. Both types of seals are also much more compression set resistant than O-rings 

and have a higher recovery rate than O-rings, allowing them to maintain constant contact with the sealing surfaces as the gap opens. 

At the forward end of the igniter, mounted to the igniter adapter, is the 

ignition initiator. The initiator utilizes the same hardware that is used in the current 

design but it is loaded with ARCADENE 360A HTPB propellant to maintain compatibility 

with the main igniter propellant. The grain design is also the same as in the current 

design since the 360A propellant is tailored to have the same burn rate as the TP1178. 

The initiator case will be insulated with Kevlar and silica-filled Hypalon to the same 

thickness that the abestos-silica NBR is applied on the current design. 

The S&A device will remain the same as that used on the current design with 

the asbestos containing parts replaced with nonasbestos materials. The S&A clutch disc 

material will be replaced with Kevlar phenolic while the S&A commutator material will 

be replaced with ceramic phenolic. 

The S&A is attached to the igniter adapter using 10 bolts with a washer and 

special packing used as an environmental seal. The S&A is sealed to the adapter using a 

radially squeezed G-T ring and a face sealing metal C-ring, the same as in the adapter to 

main SRM case. The dual O-ring seals on the AhM main rotor will remain the same as 

will the seals on the NSIs. For the external environmental seals, either an O-ring type 

seal or a formed in-place gasket material (Le., RTV) can be used on the igniter adapter 

to SRM case and the S&A to igniter adapter. 

Total weight savings for the improved versus the production igniter are 

approximately I IO pounds. The reduced grain length (32.8 inches vs. 21.6 inches) and the 

thinner insulation on igniter case account for this weight reduction. 

2.6.4 IGNITER PROPELLANT 

The igniter propellant selected for the improved igniter is ARCADENE 360A, 

which is detailed in Table 2.6.1. ARCADENE 360A is an 88 percent solids-loaded HTPB 

2-119 
TABLE 2.6.1. ARCADENE 360A. 

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INGREDIENT 

R-45 HT BINDER 

DOA 

A1 POWDER 

Fe203 

AP (60/40 200g/MA) 

TOTAL SOLIDS 

I O sps 

DENSITY 

EQUILIBRIUM Tc 

C* 

E 

GAMMA 

BURNING RATE (1000 PSI) 

PRESSURE EXPONENT 

- WT % 

10.0 

2.0 

18.0 

1.5 

68.5 

100.0 

- 

88% 

260.7 lbf-sec/lbm 

0.0655 lb/cu-in 

3508" K 

5123 ft/sec 

10.74 

1.166 

0.70 

0.48 

MECHANICAL PROPERTIES (2 in/min x head) 

- 70°F - -40°F +160 O F 

MAX. STRESS (PSI) 201 514 128 

% 30 43 26 STRAIN AT MAX. STRESS 

TANGENT MODULUS (PSI) 1540 12,100 1010 

I 

S 2-120 
propellant with a bimodal blend of ammonium perchlorate (AP) and 18 percent aluminum. This propellant is a variant of the MLRS propellant of which ARC has loaded over 

28 million pounds into MLRS motors. The higher burn rate necessary for the SRM igniter 

application is achieved by increasing the percentage of iron oxide (-1 percent vs. 

1.5 percent) and by varying the percentage of fine versus course AP in the bimodal 

blend. This propellant is completely characterized for use through the MLRS Program 

and represents very low risk in the SRM igniter application. Prior to casting the propellant, the insulated case will be barrier coated with EA-946 and then lined with ARL-151 

liner. Both of these materials have been well characterized for use with this propellant 

in the MLRS Program. 

2.6.5 IGNITER GRAIN DESIGN 

The igniter grain is shown in Figure 2.6.3 and described in Table 2.6.2. The 

grain design is a 40-point star design with the web between star points varying from 

0.20 inch at the head end to 0.05 inch at the aft end. The star tip radii are 4.90 inches 

from the igniter centerline. The maximum nominal mass flow rate is approximately 

65 percent of the DM-1 igniter firing. The molded cellulose phenolic nozzle has a throat 

diameter of 6.60 inches and will have a projected 0.030-inch total erosion on the 

diameter during the igniter firing. Maximum expected operating pressure (MEOP, +3a 

condition) is projected to be 1,660 psia. 

This grain design matches very closely the performance from the current 

production igniter. Igniter MEOP and mass flow rates compare very favorably with SRM 

firings QM-1 and QM-2 as shown in Figure 2.6.4. Igniter plume contact with the SRM 

will therefore match the current production igniter and ignition times for the SRM should 

remain the same. 

2.6.6 IGNITER INSULATION 

The improved igniter insulation selected for replacing the current asbestos 

and silica-filled NBR is Kevlar and silica-filled Hypalon. This selection comes from the 

extensive trade studies documented in Section 2.4. Preliminary estimates of insulation 

requirements, based upon reported Material Affected Rates (MARS) for the igniter in 

QM-2, using a 2.0 MAR safety factor, and reducing the igniter length results in a total 

igniter assembly insulation weight reduction of approximately 78 pounds. On the 

2-121 
R 4.90 

R 4.90 

FIGURE 2.6.3. IMPROVED SRM IGNITER GRAIN. 

2-122 
TABLE 2.6.2. SRM IGNITION SYSTEM. 

PROPELLANT 

- HTPB/AP/Al 

= 0.70 ips @ 60°F - rlooo 

GRAIN CONFIGURATION 

- 119 lb. GRAIN WEIGHT 

- 40-POINT STAR, 21.6 IN. LONG BY 16.4 IN. OD 

- 4.90 IN. RADIUS TO STAR TIPS 

- 0.20 IN. TO 0.05 IN. WALL WEB TAPER FORWARD TO AFT 

NOZZLE 

- CELLULOSE PHENOLIC 

- 6.60 IN. THROAT DIAMETER 

0.030 IN. TOTAL PREDICTED EROSION ON DIAMETER 

PERFORMANCE 

- 323 lbm/s MAXIMUM NOMINAL MASS FLOW RATE 

- 1660 PSIA MAXIMUM EXPECTED OPERATING PRESSURE (9O"F, +3a CONDITIONS) 

2-123 
ORIGINAI; PAGE IS 

OF POOR QUALITY 

u- 

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0 

( I ---__ - ----- - 7 .l-- IO 0.130 0.200 0.300 0.400 3.>00 0.600 O.?OO 0.tJOC 0.130 1.050 

TIM1 . :Lc 1.135 

FIGURE 2.6.4. IMPROVED IGNITER PRESSURE TRACE 

SRM IGNITER NOMINAL PERFORMANCE @ 60°F. 

2-124 
average, external igniter case insulation is reduced in thickness by 20 percent while the 

internal case insulation is reduced by 33 percent. The igniter initiator insulation thickness was not changed. 

The silica phenolic throat insert will be replaced with a lower cost, higher 

performance cellulose phenolic. Estimated cost savings of the cellulose versus the silica 

phenolic are approximately 85 percent for the raw material. The cellulose phenolic is 

also predicted to have 80 percent of the erosion experienced in the silica phenolic and 15 

percent less weight due to a lower density. 

These insulation thicknesses keep the igniter hardware below the required 

300'F during SRM action time and 400'F following SRM web burnout per specification 

CPW 1-3300. 

2.6.7 IGNITER SEALS 

The primary objective of this design study is to reduce potential exhaust gas 

leak paths from the igniter assembly. To achieve that goal, various hardware configuration concepts were examined to reduce the overall number of primary and secondary gas 

seals. In the current production igniter, there are 6 primary, 42 secondary, and 88 environmental seals as detailed in Table 2.6.3. In the proposed design, the case and adapter 

will be manufactured from 200 maraging steel and welded together, eliminating all caseto-adapter seals in the production design (Figure 2.6.2). Using this design, the total 

number of seals is reduced to 5 primary, 5 secondary, and 52 environmental. This provides a delta of I primary seal and 37 secondary seals. The secondary seals are drastically reduced due to the elimination of the bolts and the special bolt packing (stato-seals) 

in the igniter case to adapter joint. 

A trade study was conducted to determine what type or types of primary and 

secondary seals to utilize in the improved igniter design. The trade study is presented in 

Table 2.6.4. The primary seals selected for use on the igniter adapter to SRM case and 

the ShA-to-igniter adapter are aerospace G-T rings, These rings are radial squeeze seals 

and are commonly used to seal hydraulic fluid up to 8,000 psi. These seals are very 

resistant to extrusion due built-in nonextrusion rings, provide a positive seal at zero or 

low pressures due to radial compression at installation, and are not subject to rolling or 

spiral failures. These G-T ring seals permit sealing with larger gaps than O-rings under 

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TABLE 2.6.3. IGNITER ASSEMBLY SEALS. 

PRESENT IGNITER IMPROVED IGNITER 

PRIMARY SECONDARY ENVIRONMENTAL PRIWRY SECONDARY ENv IRONMENTAL 

S&A 

BARRIER-BOOSTER 

SHAFT 

' NSIs 

SA TO IGNITER ADAPTER 

IGNITER ADAPTER? TO 

IGNITER CASE 

IGNITER ADAPTER TO 

SRM CASE 

TOTALS 

1 1 0 1 1 

2 2 0 2 2 

1 1 11 1 1 

1 37 37 - ~ - 

0 

0 

11 

6 42 88 5 5 52 

36 Of THESE 37 SEALS FUNCTION AS SECONDARY AND ENVIRONMENTAL SEAL 

tt INCLUDES 36 SEALS LISTED AS SECONDARY * DOES NOT INCLUDE PRESSURE 

RESILIENCE (SPRINGBACK) 

COnPRf SSION SET 

RES1 STANCE 

TOUGHNESS 

(DAMAGE TOLERANCE) 

RELAXATION MODULES 

HIGH TEMPERATURE TOLERANCE 

GAP SEALING CAPABILITY 

WTIMG SURFACE 

FI NI SH/K4CHI NI NG 

EXTRUS ION RES I STANCE 

TOTALS 

TRANSDUCER SEALS 

TABLE 2.6.4. SEAL TRADE STUDY. 

O-RING STATIC FLEX LATHE 

C-RING, FACE METAL 6-1 CUT GASK- 

- RUBBER --- METAL METAL SEAL GASKET SEAL RING 0-SEAL 

7 

5 

7 

5 

4 

5 

8 

5 

46 

- 

8 10 

10 10 

4 5 

8 8 

10 10 

7 10 

6 6 

10 

63 69 

- 10 - 

RATED 1 THROUGH 10, WITH 10 HIGHEST 

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expansion of a pressure vessel as it is pressurized. Aerospace C-T rings are also designed 

to fit any groove defined in specification MIL-C-5514F. Rubber compounds allow temperature coverage from -70'F to +450°F. 

The secondary seals for the igniter adapter to SRM case and the S&A-toigniter adapter are resilient metal C-rings. These rings will seal up to 9,800 psi in a 

gland with a 32 RMS finish and can handle temperatures from cryogenic to 2,200°F. A 

metal C-ring seals at low and no pressure due to compression from the flange joint. 

System pressure then supplements the sealing force by forcing the walls of the ring 

against its mating surfaces. 

Both the G-T ring and metal C-ring seals are more compression set resistant 

than standard O-rings due to their basic designs. In dynamic loading situations where the 

gap between mating sealing surfaces tends to open, their resiliency assures that they will 

maintain contact with the sealing surfaces. By separating the seals into radial and face 

seals with different temperature capabilities for each, we are assured that no credible 

single event can cause a failure of both the primary and the secondary seal. 

The aft closure will be sealed with a single static face seal that is similar in 

design to the C-T rings described above. It consists of an "L" shaped elastomeric sealing 

element and a mating nonextrusion ring. At low or zero pressure, the static face seal 

(SFS) seals like an O-ring. A pressure increase causes the elastomer to seal more tightly 

while the nonextrusion ring precludes extrusion. SFSs can seal with clearance gaps up to 

0.015 inch and to pressures exceeding 10,000 psi. A single seal is utilized due to the fact 

that any leakage here would be into the SRM main chamber and should not compromise 

ignition (assuming leakage is not gross). 

2.6.8 SUMMARY 

The proposed igniter design utilizes a one-piece case/igniter adapter made 

from 200 maraging steel that is insulated internally and externally with Kevlar and 

silica-filled Hypalon. An aft closure bolts to the case and allows grain casting and mandrel pulling from the aft end of the igniter. An HTPB propellant, ARCADENE 360A, a 

variant on our well characterized MLRS propellant, will be utilized for the propellant 

grain. A combination of elastomeric and metal seal rings will be utilized to provide the 

minimum number of primary and secondary seals while providing superior sealing under 

all operating conditions of the SRMs. 

2-127 
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3.0 PERFORMANCE 

The selection of the heads-up trajectory as the baseline for the Block I1 SRM 

configuration results in an increase of the head-end maximum expected ,operating pressure (MEOP) to 1,085 psia from 1,015 psia. The increase in pressure was required to 

produce the higher thrust level dictated by the heads-up trajectory and is discussed in 

detail in Section 3.1. The higher MEOP requires increasing the basic case wall thickness, 

resulting in an increase of motor case weight. The 6.9 percent increase in MEOP requires an additional 0.032 inch of basic case wall. 

The increase in case wall thickness is more than offset by the reduced insulation thickness of the Block I1 asbestos-free configuration, In fact, the average 0.054-inch 

reduction of insulation thickness allows an average net 0.022-inch increase of the propellant grains' outer diameter. Table 3.0.1 presents the change in propellant and inert 

weights for the heads-up Block I1 SRM compared to the current high performance motor 

(HPM) SRM. 

The impact of changes in inert weight and propellant weight on payload 

capability are given by the following partials: 

Inert Weight: 0.0182 lbm/lbm 

Propellant Weight: 0.083 lbm/lbm 

Therefore, the 2,727-lbm increase in SRM inert weight will cause a 496-lbm reduction in 

payload capability. Also, the propellant weight reduction of 1,669 lbm will cause a 

payload capability reduction of 139 lbm. Therefore, the total change in payload capability resulting from the Block 11 improvements is a reduction of 635 Ibm. 

It must be emphasized that the 635-lbm decrease does not include the basic 

advantage of flying the heads-up trajectory rather than the current one. While ARC has 

not been supplied with a value, it is likely that the payload increase associated with the 

heads-up trajectory will more than offset the 635-lbm reduction. 

3- 1 
TABLE 3.0.1. HEADS-UP BLOCK I1 WEIGHT SUMMARY. 

LOWER DENSITY ASBESTOS-FREE INSULATION - 4,333 

IMPROVED LARC FIELD JOINT (NO FACTORY JOINT) + 260 

CASE THICKNESS INCREASE OF 0.032 INCH + 6,800 

NET SRM INERT WEIGHT CHANGE + 2,727 

PROPELLANT WEIGHT LOST TO IMPROVED JOINT CONFIGURATION - 2,560 

PROPELLANT WEIGHT GAINED ON SEGMENT ODS + 891 

(+ 0.054 - 0.032 = + 0.022 INCH) 

NET SRM PROPELLANT WEIGHT CHANGE - 1,669 

3-2 
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3.1 MOTOR BALLISTIC PERFORMANCE 

Included as one of the design study areas was an investigation into the SRM 

modifications necessary to produce the heads-up thrust history. The required nominal 

burning rate thrust history bandwidth at 60°F was presented in Enclosure 22 of a letter 

from Larry Wear, "Responses to Block I1 SRM Requests." Simply stated, the heads-up 

thrust history requires a 10-percent increase in thrust level with a reduction in burning 

time to produce the total impulse required by Specification No. CPW 1-3300. Figure 3.1.1 

presents the required thrust band presented in CPW 1-3300, while Figure 3.1.2 presents 

the heads-up thrust requirement. 

Before heads-up modifications were investigated, a simple ballistic prediction 

model was established for the current HPM. Figure 3.1.3 presents the baseline predicted 

thrust, which agrees well with the required bandwidth. Figure 3.1.4 compares the predicted pressure history with the nominal HPM curve. This simple prediction model does 

not include erosive burning behavior. This accounts for the low predicted pressure over 

the initial I5 seconds and the higher level during the 85- to 105-second time span. The 

predicted delivered vacuum specific impulse is 267.19 Ibf-s/lbm. 

Four different design parameters were investigated in an effort to produce 

the heads-up thrust history with minimum impact on the existing design. Combinations 

of propellant burning rate, propellant formulation, nozzle throat geometry, and propellant burning surface area versus web distance were studied. A number of different 

approaches were identified to produce the heads-up thrust history. Each approach also 

has a negative impact to some extent. For each approach, the impact of payload was 

calculated, and SRM components were identified that would be impacted by the approach. Table 3.1.1 summarizes the impact of the approaches that will be discussed. 

The first approach investigated was to merely increase the burning rate of 

the current propellant. This can be accomplished by adjusting the amount of iron oxide 

(FeZ03) and/or the ground-to-unground ratio of ammonium perchlorate (AP). The required base burning rate at 625 psia and 60°F was determined to be 0.386 in/s compared 

to the current 0,362 in/s baseline. Figure 3.1.5 presents the predicted thrust history 

based on a burning rate of 0.386 in/s. 

3-3 
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UPPER BOUND 

I I 1 

O+ (I 20 40 60 

TIME (set) 

1 ... 1 I 1 

100 120 1 80 10 

FIGURE 3.1.1. SPECIFICATION NO. CPW1-3300 REQUIRED THRUST BAND. 

3-4 
C 

C 

C 

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C 

nC 

GC 

'om 

0 

0 

0 

-0 

0 

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- 

CnN 

3 

CY 

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0 

0 

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0' 3 

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3 

FIGURE 3.1.2. REQUIRED "HEADS-UP TRAJECTORY" THRUST BAND. 

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0 

0 

01 

v I I 1 I I I 

in REQUIRED BAND 1 

I 

-- I 1 1 1 1 

20 40 60 80 100 120 1 o+ 0 

TIME (set) 

FIGURE 3.1.3. PREDICTED NOMINAL HPM THRUST HISTORY AT 60°F. 

3-6 
1 .m 

ACTUAL NOMINAL 

__ - PREDICTED 

NOM I NAL 

__ - PREDICTED , NOM I NAL 0 

800- 

a 600- 

W 

U 

3 

v) 

U 

," 400- 

2w0 lllr~~

0 10 20 30 40 50 60 70 80 90 100 lr 120 1 3 

FIGURE 3.1.4. PREDICTED NOMINAL HPM THRUST PRESSURE HISTORY AT 60°F. 

3-7 
TABLE 3.1.1. SUMMARY OF HEADS-UP THRUST MOD IF ICAT IONS. 

APP R 11 AC H 

1 2 3 4 

KEQUIHED BUKNING RATE, in/s 0.3&b O.3Y9 0.362 0.378 

NULLE UIAMETEH. in 53.86 56.49 51 .OU 53.86 

MEUP, psia 1,120 1.015 1,215 1,120 

PAYLOAD IMPACT U/DbAC CA5E -1,8UU -1,56U 

THICKNESS INCREASE. Ibm 

PAYLUAD IMPACT W/HIGHEH U -1,560 

STRENGTH STEEL (d CUHRENT 

THICKNESS. Ibm 

- 881 

t2.576 

- 40 

+I ,840 

COMPONENT DEVELOPMENT/ NUNE NUZZLE STKUCTUKE 2U% MEOP INCHEASE HT PROPELLANT 

WEDESIGN REl)UlRED* ANU FLEX-BEARING LIKELY HIGH ENOUGH DEVELOPMENT 

TO HEQUIKE COMPLETE 

REDESIGN 

ASSUMES CUKKENT PBAN CAN BE TAILORED TO REQUIRED RATE. 

3-8 
1 

I 

I 

60 80 

TlMF I,,,\ 

100 

REQUIRED BAND 

PREDICTED 

120 1 

I 

I 

I 

FIGURE 3.1.5. PREDICTED HEADS-UP THRUST HISTORY AT 60°F AND rb = 0.386 IN/S. 

3-9 
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While the required thrust band is achieved, this approach would increase 

maximum expected operating pressure (MEOP) to approximately 1,120 psia. The increase 

in MEOP would prohibit use of existing hardware. The required increase in case wall 

thickness using D6AC steel would result in an approximate I ,800-lbm payload capability 

loss. This approach will involve no additional redesign or development beyond the concept currently proposed for the Block I1 SRM. 

The second approach also involved an increase in burning rate, but nozzle 

throat area was also increased to retain MEOP at its current level. The required burning 

rate and initial throat diameter were determined to be 0.399 in/s and 56.490 inches, 

respectively. Other minor nozzle modifications would include reoptimization of the 

nozzle exit cone contour and use of a carbon phenolic material with a 15 percent lower 

nozzle erosion rate. 

While satisfying, in general, the required thrust band, the resulting delivered 

specific impulse is reduced to 265.24 Ibf-s/lbm from the current 267.19. The loss of 1.95 

Ibf-s/lbm delivered specific impulse results in a payload reduction of 1,560 Ibm. Also, 

the increase in nozzle throat diameter is sufficiently large to require the redesign of the 

metal backup structure of the nozzle. 

The third approach investigated was a nozzle throat diameter reduction. Like 

the first approach, this method will increase MEOP. However, the nozzle diameter 

reduction will increase the delivered specific impulse. Also, the reduction will structurally allow the use of carbon/carbon as the nozzle throat material (lower erosion rate) 

within the envelope of the existing metal parts. The nozzle diameter required for this 

approach is 51.0 inches. 

Again, the desired thrust band is achieved. MEOP is 1,215 psia, but the 

delivered specific impulse is increased to 270.41 Ibf-s/lbm. The delivered I SP and increased inert weight trade to yield a S87-lbm decrease in payload. It is likely that such a 

great increase in MEOP will cause other components besides the case wall to increase in 

weight. Therefore, the payload capability would likely decrease somewhat below the 

number just presented. 

The fourth approach considered was the use of a higher energy and density 

HTPB propellant. The HTPB propellant will also likely have a pressure exponent that is 

3-10 
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somewhat higher than that of the PBAN propellant. This approach will require tailoring 

of the burning surface area versus web history such that the maximum burning surface 

area and the minimum surface area are closer to the web averaged mean. The required 

burning rate for the HTPB propellant would be 0.378 in/s at 625 psia and 60'F. 

Again, this option results in raising MEOP to approximately 1,120 psia. The 

increased case wall thickness would result in a net 40-lbm payload increase. The increased wall thickness is offset by a 0.6-lbf-s/lbm increase in theoretical specific impulse and 20,800 Ibm of extra propellant due to the higher propellant density. Also 

included is a 2,000-lbm increase in the insulation required for HTPB propellant use. 

In addition to the four specific options discussed, a great number of other 

specific combinations of the four parameters exist. It is clear, however, that the current 

nozzle flex-bearing support structure cannot accommodate the increase in nozzle throat 

diameter (56.49 inches) required to maintain MEOP at its current level. Therefore, if the 

Block I1 SRM is to produce the heads-up thrust history, some increase in MEOP will 

result. 

Based on the Block I1 SRM design studies to date, the best way to achieve the 

heads-up trajectory is the first approach (raise burning rate and maintain nozzle throat 

diameter). This is largely based on the minimum amount of additional development or 

redesign associated with this approach. Also, several factors can reduce the projected 

MEOP (at 90'F) increase from 1,015 psia to 1,120 psia associated with the selected 

opt ion. 

The MEOP of the current HPM configuration is driven by the erosive burning 

pressure spike that occurs on ignition. The pressure overshoot decays exponentially over 

the first 10 seconds of burning. The observed maximum pressure of the current HPM 

SRM is 30 to 40 psia higher than the maximum pressure that would result if there were 

no erosive burning. Therefore, reducing the burning surface area over the initial 10 

seconds to compensate for the erosive burning will allow an MEOP reduction of approximately 35 psia. Applying this reduction to the earlier stated MEOP of the selected 

heads-up design approach (increase burning rate) reduces the MEOP of the proposed 

heads-up trajectory configuration to 1,085 psia. This is a 6.9-percent increase over the 

1,015 psia MEOP of the current HPM SRM. 

3-11 
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The reduction of the initial burning surface area during the first 10 seconds is 

achieved by modifying the inhibitor pattern on the aft face of the two center segments. 

The amount of inhibited surface required to reduce initial pressure by 35 psia is approximately 10,000 in of burning surface area. The modified inhibitor pattern is shown in 

Figure 3.1.6. The width of the inhibitor projections is 8 inches, which is equal to twice 

the web distance burned at 10 seconds. Over the burning period of 10 to 20 seconds, the 

burning surface area closely matches that of the current grain configuration. Figure 

3.1.7 compares the total burning surface area produced by the original and modified grain 

face inhibitor configurations. 

2 

The predicted ballistic performance produced by the heads-up modifications 

(burning rate and grain inhibition) is presented in Table 3.1.2. 

3-12 
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3.2 JOINTS AND SEALS 

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The heads up mission internal MEOP has been determined to be 1,073 psig at 

the forward end of the SRM. Following the calculations of Section 2.3.1.5, the minimum 

membrane wall will increase 

= 0.491 inch. 

The corresponding nominal membrane wall thickness will be 0.51 1 inch. 

The impact of the higher operating pressure on the metal case joint is expected to be minimal. The critical safety factors per Section 2.3 were large enough to 

accommodate the 1.069 increase in loading. Therefore, no change in metal thickness or 

bolting is anticipated. Similarly, no change in the joint insulation arrangement or thickness will be required. 

The primary impact on the nozzle-to-case joint for the heads-up trajectory is 

due to the revised loads caused by motor design changes made to meet the new performance requirements. Chamber pressure in the joint region increases by approximately 

6.9 percent to 972 psig. The nozzle blow-off load actually decreases slightly to 1.14 x 

10 pounds. The reduction is due to a 10-percent increase in thrust, which offsets the 

pressure load. Joint deflections and stresses are a function of both the reduced blow-off 

load, which is reacted by the bolted interface, and the increased pressure acting on the 

motor aft dome and nozzle fixed housing. A comprehensive analysis is required to evaluate accurately the impact of this combination. However, since a complete reanalysis of 

the structure was not possible, a conservative approach was used that assumed all loads 

increased by 6.9 percent. A review of the margins of safety previously shown in Figures 

2.3.31 and 2.3.38 for the joint hardware and insulation, respectively, shows that all 

critical margins exceed 0.069 and, therefore, the current design has adequate strength to 

carry the increased load. The revised minimum margins would be +0.09 for the hardware 

and +0.46 for the insulation. To evaluate the seal function, the gap versus pressure data 

reported in Section 2.3 was extended to 972 psig. This is included as Figure 3.2.1, which 

shows the maximum gap at the primary seal would be 0.0063. When multiplied by two 

and compared to the conservative "springback" of 0.012 inch recommended by the seal 

manufacturer, a negative margin of safety results, indicating that some redesign is 

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0 100 200 300 400 500 600 700 800 900 1000 1100 

PRESSURE, PSI 

FIGURE 3.2.1. SEAL GAPS VERSUS EXTENDED MOTOR PRESSURE. 

3-17 
required. The design changes that could be investigated would include increasing the bolt 

pre-load or size, increasing the primary seal size, and increasing the nozzle flange thickness. It is not possible to determine which of these approaches would provide the optimum design; however, we believe the weight impact would certainly be less than that 

incurred in going from the 51-L design to the current design for the HPM. 

3-18 
3.3 ASBESTOS-FREE INSULATION 

The thermal analyses described in Section 2.4 used the current HPM ballistics 

in determining boundary conditions and burn times. This approach was necessary so that 

a direct comparison to the available data and predictions for the HPM design with asbestos insulation could be made. To assess the impact of the heads-up trajectory on the 

nonasbestos insulation, select analyses were rerun using the appropriate ballistics. The 

higher operating pressure for this heads-up design will increase the amount of insulation 

degradation while the shortened burn time will decrease insulation degradation. The 

shortened burntime dominates and the net result is analytically predicted to be a 

4.3-percent decrease in degraded insulation thickness. 

An assessment of the thermal impact of the heads-up trajectory was also 

performed on the nozzle throat. Ablation analyses were performed at the throat location 

to predict the erosion, char, and resulting safety factors for both the current HPM and 

the heads-up ballistics. The results show that the heads-up trajectory will have a 

14.4-percent decrease in throat erosion and a 3.8-percent increase in char thickness. 

These predictions combine to yield a IO-percent increase in the safety factor at the 

throat for the heads-up design when compared to the HPM design. 

3-19 
3.4 NOZZLE DESIGN 

The motor design changes required to meet the heads-up trajectory performance requirements impact the nozzle design in terms of both thermal response and 

structural integrity. The effects of increased mass flow and revised burntime on liner 

erosion and insulation requirements are discussed in Section 3.3 of this report. Similarly, 

the nozzle-to-case joint region is considered in Section 3.2. The remaining nozzle structural impacts resulting from motor changes are discussed below. 

The increases in MEOP and thrust required to meet the heads-up requirements result in revised loading on the nozzle shell structure and on the flex bearing. 

Since the basic nozzle concept was retained in the Block 11 SRM design, no significant 

stress analysis was conducted as part of this program. Attempts to obtain the documented stress analysis for the current design were unsuccessful. Therefore, only a 

qualitative evaluation of design impacts due to this load increase was possible. In those 

places in the nozzle shell that are critical for internal pressure, either motor pressure or 

the correspondingly increased nozzle internal pressure profile, the stresses will be increased by approximately 6.9 percent. Discussions with NASA personnel indicate that 

some areas of the nozzle are not adequate for this increase; therefore, increases in 

material thicknesses would be required. The weight impact of this increase is expected 

to be less than 6.9 percent of the nozzle structural weight, since not all areas require 

increases. 

In those areas of the nozzle assembly critical for blow-off loads, including the 

flex bearing, loads will vary somewhat less than 6.9 percent. This is because the thrust 

increases more than MEOP; thus, blow-off loads that are pressure loads minus thrust load 

increase less than pressure loading by itself. In addition to reviewing structural adequacy 

for this load, a clearance check for deformed structure should be conducted. In summary, the changes in nozzle structure are expected to include only local thickness increases with the resulting weight increase. 

3-20 
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3.5 IGNITER 

The improved Block I1 SRM igniter performance was discussed in 

Section 2.6.5. The igniter was designed to duplicate the performance of the current 

production igniter, thereby having no impact on the overall motor performance. The 

igniter propellant, discussed in Section 2.6.4, is a well characterized version of 

ARCADENE 360A, an HTPB, 88 percent solids-loaded composite propellant. This propellant is a variant of the MLRS propellant of which ARC has loaded over 28 million pounds 

into MLRS motors. 

Based on the above, the ballistic performance of the igniter represents very 

low risk to the SRM Block I1 Program. 

3-21 
4.0 PRELIMINARY 

4.1 DESCRIPTION 

DEVELOPMENT AND VERIFICATION 

The Space Shuttle Block I1 SRM is used as a major 

PLAN 

subsystem of the solid 

rocket booster (SRB) for the shuttle vehicle. The shuttle vehicle booster will consist of 

two SRBs, each of which will use a solid rocket motor. The two SRMs will operate in 

parallel with the main engines, provide impulse, and enable thrust vector control to 

propel and control the space shuttle vehicle from lift-off to SRB staging. The SRM will 

consist of a lined, insulated, segmented rocket motor case loaded with solid propellant; 

an ignition system complete with an electromechanical safe and arm device, initiators, 

and igniter assembly; a moveable nozzle; an exit cone with a linear-shaped charge for 

severance and a nozzle plug; instrumentation; and integration hardware including electrical brackets, systems tunnel base, grounding provisions, stiffener rings, and attachment provisions to the forward and aft SRB skirts and the SRB/external tank (ET) 

attachment ring. 

The four SRM casting segments are comprised of six case segments. The 

bolted joint configuration will be integrated into all case segment field joints. The forward and aft closure and SRM skirt attachment joints will retain the existing tang and 

clevis joint concept. Closures will be assembled to case segments with factory joints. 

The Block I1 case segments, roll formed from MAC steel into case lengths equal to 

casting segments, will retain the weld-free concept of the Block I motors. The same 

TPH-1148 propellant will be used. However, design improvements to the insulation/ 

inhibitor system at the field case joints will be incorporated. All materials will be 

asbestos-free. 

The ARC Block 11 SRM will be a highly reliable motor that permits expeditious assembly and maximum visibility 

launch facility. 

4.2 OBJECTIVE AND SUMMARY 

This Program Plan describes 

of joint integrity during SRM stacking at the 

how ARC proposes to develop and verify the 

Block I1 SRM using the NASALaRC/ARC-designed bolted joint protected by an unvented 

4-1 
insulation joint. The program milestone schedule shown in Figure 4.2.1 highlights major 

events that will be accomplished and documented. 

This Program Plan describes ARC'S understanding of the effort required to 

develop and verify the Block I1 SRM. ARC has the nucleus of the management and 

technical disciplines to perform the Block I1 SRM Development and Verification (D&V) 

Program successfully. An aggressive staffing plan will be developed and initiated in 

advance of Block I1 SRM Program initiation. 

Block I1 SRM requirements will be verified by analysis, inspection, and testing 

from the component to the subscale to the system level. ARC understands SRM D&V 

requirements and offers the program flow logic diagrammed in Figure 4.2.2 to accomplish NASA's objectives in a timely manner. The D&V matrix shown in Figure 4.2.3 

displays the analysis and tests planned to verify each component. 

4.3 MANAGEMENT 

Overall responsibility for the Block I1 SRM rests with the SRM Project Director, who reports directly to the Senior Vice President and General Manager of the ARC 

Propulsion Division. The SRM Project Office at our Camden, Arkansas facility will 

direct all D&V tasks, including the detailed motor and facilities design effort. Initial 

propellant/liner/insulation laboratory testing will be accomplished at our Gainesville, 

Virginia facility, and the technology and/or personnel will be transferred to Camden to 

assure transition from the laboratory to full-scale operations. Figure 4.3.1 shows the 

SRM project organization structure and responsibilities. 

4.4 SRM DESIGN AND ANALYSIS 

The ARC Block I1 Conceptual Design Motor eliminates all deficiencies identified in the Shuttle 51-L SRM. The integrity and high reliability of the ARC Block I1 SRM 

will be verified during the planned program. This design offers marked advantages with 

regard to ease of assembly and inherent reliability as compared with the SRM presently 

being designed and tested under the SRM Recovery Program. The design trades conducted in reaching the SRM design recommended by ARC were heavily weighted toward 

flight safety and reliability and resulted in the selection of materials, components, and 

designs having an existing database. The two novel concepts proposed by ARC -- bolted 

4-2 
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SRM OVERALL PROJECT MANAGEMENT 

PROJECT CONTRACT MANAGEMENT 

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RELIABILITY 

SYSTEMS SAFETY 

SYSTEMS REQ'T 

ANALY SI S 

CONFIGURATION 

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LOGISTICS SUPPORT 

TRANSPORTATION 

ANALY SI S 

BUSINESS MANAGEMENT 

FIELD SUPPORT 

SAFETY 

MKHANICAL DESIGN 

SYSTEMS PRODUCT 

ENGINEERING ENGINEERING MANUFACTURING 

BALLISTICS DESIGN 6 

STRUCTURAL DESIGN 6 

ANALYSIS 

QUAL I TY SUBCONTRACTS 6 

TESTING ASSURANCE MATERIAL 

ANALYSIS 

THERMAL DESIGN 6 

ANALYSIS 

TVC DESIGN 

GAS DYNAMIC 

ANALYSIS 

PROCESS ENGINEERING 

TOOLING ENGINEERING 

* MANAGEMENT 

MFG PLANNING TEST ENGINEERING QUALITY ENGR 

MFG SUPPORT TESTING PROCESS 6 

MOTOR MFG DATA REDUCTION PERSONNEL 

MOTOR ASSY 

MOTOR PACK/SHIP INSPECTION 

MOTOR REFURB PRODUCT DATA 

FACILITY ENGR REVIEW 

READINESS REVIEWS 

PERSONNEL TRAINING 

C ER T I F I C AT IO ti 

FIGURE 4.3.1. SRM PROJECT ORGANIZATION. 

4-6 
case field joint and double-length case segments -- are based on sound engineering technology and will be thoroughly verified during the D&V Program. The bolted case field 

joint is based on extensive work done by the NASA Langley Research Center and confirmed by ARC analysis. Bolted, face-sealed joints are typically used for large diameter 

nozzle-to-case joints. The double-length case segment concept has been studied by 

Ladish Co., Inc. and rated feasible with additional foundry, forging, and heat treatment 

facilitization. ARC has extensive experience in developing asbestos-free insulation; 

several rocket systems are in production using this improved material. These concepts 

have been analyzed to define their structural, thermal, and gas dynamic characteristics. 

ARC plans to work closely with the MSFC Aeroballistic lab trajectory group to assure 

that ARC'S SRM design satisfactorily interfaces with the shuttle vehicle's flight performance. 

The Block I1 Engineering Task will start with the review and design impact 

analysis of all SRM system and performance requirements. The preliminary Contract 

End Item (CEO Specification (Part I) submitted with this report will be compared to 

overall SRM and mission requirements following ATP to establish a baseline. We anticipate no problem in meeting the SRM requirements; however, our program planning is so 

structured that any problem that arises will receive immediate attention and action. 

4.5 COMPONENT AND SUBSCALE TESTS 

Component tests were selected to provide empirical data for analysis, to 

validate analytical models, and to demonstrate the feasibility of ARC Block I1 SRM 

design concepts. Table 4.5. I summarizes the component tests proposed. Parallel with 

the early case subcontractor effort, ARC will conduct component development of asbestos-free insulation, insulation joints, seals, bond characterizations, and the igniter. ARC 

plans aeroheat testing exposure of systems tunnel sections at ARC. ARC offers its free 

jet testing facility, which will duplicate SRM flight conditions to provide early empirical 

data to verify the systems tunnel design. Some special tests, such as the case material 

properties evaluation, have previously been performed by MTI yielding a comprehensive 

database. 

4.6 COMPONENT AND MOTOR MANUFACTURING 

Full-size motor manufacturing will begin with cases fabricated from current 

length forgings and will use bolted joints. ARC has selected this approach for 

4-7 
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CAS E MATERIAL 

- 

PROPERTIES 

FABRICATION 

RING 

TABLE 4.5.1. COMPONENT TESTS. 

OBJECTIVE TEST QTY 

CHARACTERIZE TBD 

MATERIAL AT 

FORCED EXTREMES 

OF SPEC I FICATION 

DEVELOP MACHINING 6 

TECHNIQUES 

MANUFACTURING VALIDATE MFG 

TOLE RAN C E TOLERANCES 

HEAT TREAT DEVELOP & 

CHARACTERIZE 

HEAT TREAT 

PROCESS 

HYDROTEST EVALUATE JOINT, 

VERIFY CASE 

INTEGRITY 

HYDROBURST VERIFY CASE 

DESIGN MARGIN 

4-8 

10 

TBD 

1 FWD, 

1 CENTER, 

1 AFT 

SECTIONS 

WITH FWD 

& AFT 

CLOSURES 

2 FWD, 

1 AFT 

TEST CONDITIONS 

STANDARD ASTM TEST 

SPECIMENS, HOT/COLD 

TEMP, SALT 

IMMERSION, SURFACE 

TREATMENT VARIED 

FOR STRESS 

CORROSION 

AMBIENT 

END RINGS 146-INCH 

END RINGS 

SUBSCALE & FULL 

SCALE RINGS 

INSTRUMENT FOR Pc, 

DEFLECTION, STRAIN 

INSTRUMENT FOR Pc, 

DEFLECTION, STRAIN 
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CASE & 

NOZZLE 

FIELD 

JOINT 

TEST 

FIELD 

HANDLING 

(1/3 SCALE 

MOTOR) 

SEAL 

ENVIRONMENTS 

(INCLUDE ALL 

SEALS) 

SEAL 

HAND I NG 

BOLT STRESS 

CORROSION 

BOLT 

PRELOAD 

TABLE 4.5.1. (CONTINUED). 

OBJECT I VE 

EXAMINE ASSEMBLY 

& DISASSEMBLY 

VERIFY INTEGRITY 

AT ENVIRONMENTAL 

AND TOLERANCE 

EXTREMES 

VERIFY SEAL 

INTEGRITY 

UNDER ADVERSE 

HANDLING 

VERIFY INTEGRITY 

UNDER LONG TERM 

EXPOSURE TO 

STRESS 

VERIFY BOLT 

PRELOAD NDT 

METHOD 

4- 9 

TEST QTY 

2 

10 EA 

6 

6 

6 

TEST CONDITIONS 

STEEL/INSULATION JOINT 

SEGMENT; AMBIENT 

ENVIRONMENTS 

METAL SECTIONS, 

BASELINE & ALTERNATE 

SEALS, TEMPERATURE & 

HUMIDITY EXTREMES, 

MAXIMUM GAPS 

CLEAN & CONTAMINATED 

SURFACES, IMPOSED 

DAMAGE, TEMPERATURE & 

HUMIDITY EXTREMES 

METAL JOINT SECTION; 

SALT WATER IMMERSION, 

70% ULT PRELOAD 

3 SETS OF ULTRASONIC 

EQUIPMENT; IMPOSED MAX 

RANGE OF CALIBRATION; 

MATCH ACTUAL VERSUS 

NDT BOLT EXTENSION 
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TABLE 4.5.1. (CONTINUED). 

OBJECTIVE TEST QTY TEST CONDITIONS 

CAS E MATER I A L CHARACTERIZE 6 EACH UNIAXIAL, BIAXIAL, 

INSULATION PROPERTIES STRUCTURAL, MATERIAL, ZERO TIME, ACCELERATED, 

LINER, & THERMAL, & BOND TEST & REAL TIME AGING 

INHIBITOR, PROPERTIES AT 

NOZZLE PROCESS & 

AB LAT I V ES ENVIRONMENT 

EXTREMES 

EROSION CHARACTERIZE 6 EACH SUBSCALE EROSION 

MOTOR ABLATIVES MATERIAL, TEST MOTORS 

TEST 

CONDITION 

FLOW EVALUATE MOTOR 3 EACH INDUCED VORTEX & 

VISUALIZATION COMPONENTS UNDER CONDITION CIRCUMFERENTIAL 

IMPOSED ADVERSE FLOW 

FLOW; VALIDATE 

FLOW MODEL 

4-10 
COMPONENT - TEST 

TABLE 4.5.1. (CONTINUED). 

OBJECTIVE TEST QTY 

IGNITER SEAL VERIFY INTEGRITY 10 EA 

ENVIRONMENTS AT ENVIRONMENT & 

TOLERANCE 

EXTREMES 

PROPELLANT, CHARACTERIZE 6 EA 

LINER, BOND 

INSULATION 

BOND 

CHARACTERIZATION 

COND IT I ON 

HARDWARE VERIFY 2 

HYDROTEST HARDWARE 

INTEGRITY 

HARDWARE VERIFY DESIGN 2 

HYDROBURST MARGIN 

DEVELOPMENT EVALUATE IGNITER 6 

STATIC TESTS BALLISTICS. 

STRUCTURE 

DEVELOPMENT EVALUATE IGNITER 2 

ENVIRONMENT PERFORMANCE 

TEST 

MAIN GRAIN EVALUATE IGNITER 2 

IMPINGEMENT FLOW FOR POTENTIAL 

TEST GRAIN DAMAGE 

4-11 

TEST CONDITIONS 

IGNITER HARDWARE & 

MOTOR INTERFACE 

SECTION; BASELINE/ 

ALTERNATE SEALS, TEMP/ 

HUMIDITY/TOLERANCE 

EXTREMES 

AMBIENT ENVIRONMENTS 

PREDICT/MEASURE 

DEFLECTION, STRAIN 

AMBIENT, MEASURE STRAIN 

OPEN AIR, 2 AMBIENT, 

2 HOT, 2 COLD, WITNESS 

PANELS, HIGH FREQ Pc 

TRANSDUCERS 

FULL SEQUENTIAL 

EXPOSURE TO ENVIRONMENTS, 

STATIC TEST 1 HOT, 1 COLD 

OPEN AIR, HIGH FREQ Pc 

INERT FORWARD SEGMENT, 

1 HOT, 1 COLD 
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Development Motor tests DM-IO, 11, and 12 to allow early full-scale testing, DM-13 and 

Qualification Test motors would use longer, weld-free case segments incorporating the 

bolted joint design. Table 4.6.1 shows the configuration, quantity, and schedule of case 

segments. Tables 4.6.2 and 4.6.3 add the planned fabrication schedule for nozzle assemblies and igniter hardware. For the total D&V Program, motor preparation, insulation, 

and lining and casting schedules are shown in Table 4.6.4. 

4.7 SYSTEM LEVEL TESTS 

System-level tests are chosen to represent the SRM as it will function under 

space shuttle mission conditions. System level D&V tests are summarized in Table 4.7.1. 

Figure 4.7.1 shows the SRM Joint Evaluation Simulators (JES) test series. 

JES will be fabricated from existing SRM standard weight hardware with the new bolted 

joints welded in place of the clevis and tang joints. Standard weight segments from the 

existing inventory will be used to provide the required case safety factor with the higher 

projected MEOP for the heads-up trajectory. The JES hardware will be loaded with inert 

propellant overlaid at the joints with a thin layer of live propellant as in the current 

Recovery Program tests. Nozzle Joint Evaluation Simulators (NJES) similar to the 

current NJES test series will be fabricated and tested at our Camden facility. Figure 

4.7.2 defines the NJES test series. 

Transient Pressure Test Articles (TPTA) similar to JES units will be fabricated and tested at NASA/MSFC with applied structural launch flight loads. An Assembly Test Article consisting of 160-inch-long bolted joint hardware will be shipped to 

NASA/Kennedy Space Center. Eight tests of assembly and disassembly methods will be 

conducted with the baseline and alternate seals. Seal verification tests of each joint will 

be conducted at 50, 100, and 200 psia. Video camera coverage will be provided from 

exterior and interior views during assembly, leak test, and disassembly. The test quantity was selected to individually test the variables shown in Figure 4.7.3. 

One set of SRM hardware including forward closure, center segments, ET 

attachment ring, aft stiffener segments, aft closure, and aft skirt will be shipped to 

NASAlMSFC for assembly and test as a Structural Test Article (STA). A full series of 

SRM launch and flight structural loads will be conducted on the STA both at nominal load 

and nominal loads multiplied by full design margins. This test will verify SRM structural 

integrity . 4-12 
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DM- 11 

DM-12 

DM- 13 

QM-9 

QM- 10 

QM- 11 

QM- 12 

TABLE 4.6.2. NOZZLE FABRICATION SCHEDULE. 

(NUMBER = MONTHS ARO) 

HARDWARE 

22N 

24N 

25N 

25N 

51N 

51N 

51N 

51N 

4-14 

DELIVER TO ARC 

32N 

34N 

36N 

36N 

60N 

60N 

60N 

60N 
HYDROBURST 

IGN-1 

IGN-2 

IGN-3 

IGN-4 

IGN-5 

IGN-6 

IGN-7 

IGN-8 

IGN-9 

IGN-10 

JES 1 

JES 2 

JES 3 

JES 4 

JES 5 

JES 6 

JES 7 

JES 8 

JES 9 

JES 10 

NJES-1 

NJES-2 

NJES-3 

NJES-4 

NJES-5 

NJES-6 

NJES-8 

TPTA- 1 

TPTA-2 

TPTA-3 

TPTA-4 

TPTA-5 

TPTA-6 

TPTA-7 

TPTA-8 

TPTA-9 

TPTA- 10 

DM- 10 

DM- 11 

DM-12 

DM-13 

QM-9 

QM- 10 

QM- 11 

QM- 12 

NJES-7 

TABLE 4.6.3. IGNITER MANUFACTURING SCHEDULE. 

(NUMBER = MONTHS ARO) 

HARDWARE 

8N 

8N 

9N 

10N 

10N 

10N 

10N 

11N 

11N 

11N 

11N 

15R (IGN-1) 

15R (IGN-2) 

15R (IGN-3) 

15R (IGN-4) 

15R (IGN-5) 

15R (IGN-6) 

15R (IGN-7) 

15R (IGN-8) 

15R (IGN-9) 

15R (IGN-10) 

20R (IGN-1) 

20R (IGN-3) 

20R (IGN-4) 

20R (IGN-6) 

20R (IGN-7) 

20R (IGN-8) 

25R (IGN-1) 

25R (IGN-3) 

25R (IGN-5) 

25R (IGN-7) 

25R (IGN-8) 

20R (IGN-2) 

20R (IGN-5) 

25R (IGN-2) 

25R (IGN-4) 

25R (IGN-6) 

25R (IGN-9) 

25R (IGN-10) 

31R (IGN-1) 

31R (IGN-2) 

31R (IGN-3) 

31R (IGN-4) 

36R (IGN-6) 

36R (IGN-7) 

36R (IGN-8) 

36R (IGN-5) 

INSULATED 

- 

10 

10 

11 

11 

11 

11 

12 

12 

12 

12 

16 

16 

16 

16 

16 

16 

16 

16 

16 

16 

21 

21 

21 

21 

21 

21 

21 

21 

26 

26 

26 

26 

26 

26 

26 

26 

26 

26 

32 

32 

32 

32 

37 

37 

37 

37 

4-15 

L I NED/ LOADED 

10 

10 

11 

11 

11 

11 

13 

13 

13 

13 

18 

18 

18 

18 

18 

18 

18 

18 

18 

18 

23 

23 

23 

23 

23 

23 

23 

23 

28 

28 

28 

28 

28 

28 

28 

28 

28 

28 

34 

34 

34 

34 

38 

38 

38 

- 

38 
JES-1 THRU 10 

NJES 1 THRU 10 

TPTA 1 THRU 4 

DM- 10 

DM- 11 

DM- 12 

DM-13 

QM-9 

QM- 10 

QM-11 

QM-12 

TABLE 4.6.4. MOTOR MANUFACTURING OPERATIONS. 

REFURB 

17 

24 

22/23 

31 

31 

31 

N 

N 

N 

N 

52 

INSULATION 

19 

26 

24 

34 

34 

34 

44 

63 

63 

63 

63 

LEGEND: lSt NUMBER = MONTH OPERATION BEGINS 

4-16 

LINING 

21 

28 

25 

36 

36 

36 

46 

64 

64 

64 

64 

CASTING 

21 

28 

25 

36 

36 

36 

46 

64 

64 

64 

64 

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22 

29 

26 

30 

39 

39 

47 

65 

65 

65 

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4-21 
Four Development Test Motors were selected to allow for the following: 

Design iteration; 

Design evaluation at 20°F, 60"F, and 90°F; 

Design evaluation after sequential exposure to SRM environments. 

Four Qualification Motors were selected to provide duplicate data at 

environmental extremes. 

4-22 
4.8 FACILITIES, EQUIPMENT, AND TRANSPORTATION 

The Development and Verification Plan for facilities, equipment, and transportation will be accomplished in two phases. Phase I will include construction of those 

facilities that are absolutely necessary for the production of a single-test SRM. Some of 

the production processes can be achieved by rescheduling production of other programs. 

In addition, temporary facilities can be used. Partial completion of certain facilities will 

also aid in the implementation of a "bare bones" operation during Phase I. Phase I1 

construction will include completion of the remaining, less essential facilities. 

By rescheduling the MLRS Program so that production is completed prior to 

the required delivery dates, the mixers, grinders, storage facilities, conditioning facilities, and labor force from the MLRS Program can be used in the production of SRM 

propellant until a full-up production line is operational. This accelerated schedule can be 

achieved by working extra days and shifts. 

Temporary equipment can be used in several areas to meet the D&V production schedule. Mobile boiler plants will provide heating for comfort, curing, and processing until the completion of Phase 11. Lift house and rail loading functions will be handled 

by mobile cranes until a permanent structure is built. 

Partial completion of the road network, steam system, and power will allow 

for early production. Sewage will be collected in collection tanks and hauled to the 

existing sewage treatment plant until the sewage system is complete. 

Phase I construction will permit production of the D&V SRM within 18 months 

of contract award. Phase I facilities will include 

Case preparation, 

Casting tooling assembly, 

Casting tooling clean-up, 

Casting pits (31, 

Casting control room, 

X-ray, 

Cast house, 

"Lift house" (temporary mobile crane), 

4-23 
Segment assembly, 

Equipment/tooling storage, 

Steam plant (mobile), 

Roads (partial network), 

Site clearing (partial), 

Sewage system (collection tasks), 

Rail loading/unloading crane (temporary mobile crane). 

Phase I1 construction will parallel the construction in Phase I. It would, 

however, have lower priority until Phase I portions are complete. Phase I1 construction is 

expected to be complete about 2 years after contract award (see Figure 4.8.1). 

Transportation of the SRM will pose no special problem. Table 4.8.1 is 

provided as a routing schedule. 

4-24 
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TABLE 4.8.1. PROPOSED RAIL ROUTING OF SRM HARDWARE. 

CAMDEN, ARKANSAS TO KENNEDY SPACE CENTER 

LOAD ON EAST CAMDEN AND HIGHLAND RAILROAD AND SWITCH TO SOUTHERN 

PACIFIC AT CAMDEN 

SOUTHERN PACIFIC TO NEW ORLEANS VIA SHREVEPORT 

SEABOARD RAILROAD TO JACKSONVILLE, FLORIDA 

FLORIDA EAST COAST RAILROAD TO VANDENBERG AIR FORCE BASE/KENNEDY 

SPACE CENTER 

I 

I 

I 

I 

CAMDEN, ARKANSAS TO WESTERN TEST RANGE 

LOAD ON EAST CAMDEN AND HIGHLAND RAILROAD AND SWITCH TO SOUTHERN 

PACIFIC AT CAMDEN 

SOUTHERN PACIFIC DIRECT INTO WESTERN TEST RESEARCH FACILITY VIA 

LAS VEGAS, NEVADA 

CAMDEN, ARKANSAS TO MORTON THIOKOL INC., CORINNE, UTAH 

LOAD ON EAST CAMDEN AND HIGHLAND RAILROAD AND SWITCH TO SOUTHERN 

PACIFIC AT CAMDEN 

SOUTHERN PACIFIC TO KANSAS CITY SOUTHERN AT SHREVEPORT, LOUISIANA 

KANSAS CITY SOUTHERN TO NORTH PLATTE, NEBRASKA VIA KANSAS CITY 

UNION PACIFIC FROM NORTH PLATTE TO CORINNE, UTAH 

MORTON THIOKOL INC., CORINNE, UTAH TO CAMDEN, ARKANSAS 

THE REVERSE OF THE ROUTING FROM CAMDEN TO MORTON THIOKOL INC. AS 

DESCRIBED ABOVE WILL BE UTILIZED. 

4-26