Shuttle Solid-Rocket Boosters

The solid-rocket motors for the Titans III and IV carried the evo­lution of solid-propellant technology from the significant achieve­ments of the Polaris and Minuteman to a new level. The next step yielded the still larger solid-rocket boosters (SRBs) on the Space Shuttles. After UTC had developed the 7-segment solid-rocket mo­tors for the Titan, NASA decided in March 1972 to use SRBs on the shuttle. Even before this decision, the Marshall Space Flight Center had provided contracts of $150,000 each to the Lockheed Propulsion Company, Thiokol, UTC, and Aerojet General to study configura­tions of such motors. Using information from these studies, NASA issued a request for proposals (RFPs) on July 16, 1973, to which all four companies responded with initial technical and cost proposals in late August 1973, followed by final versions on October 15.

Because the booster cases would be recoverable, unlike those for the Titan III, and because they had to be rated to carry astronauts, they needed to be sturdier than their predecessors. Lockheed, UTC, and Thiokol all proposed segmented cases without welding. Al­though Aerojet had been an early developer of such cases, it ignored a requirement in the RFP and proposed a welded case without seg­mentation, arguing that such a case would be lighter, less costly, and safer, with transportation by barge to launch sites from Aero-

Space Shuttle solid-rocket booster in a test stand at a Thiokol test site in 1979. (Photo courtesy of NASA)

jet’s production site. Had Aerojet won the contract, it is possible that the Challenger disaster never would have occurred. However, the source evaluation board with representatives from five NASA centers and the three military services ranked Aerojet last, with a score of 655 for mission suitability. By contrast, respective scores for Lockheed, Thiokol, and UTC were 714, 710, and 710. The board selected Thiokol as winner of the competition, based on its cost, the lowest of the three, and also its perceived managerial strengths. NASA announced the selection on November 20, 1973.19

Since Thiokol had plants in Utah, NASA administrator James C. Fletcher’s home state, the decision was controversial. Lockheed pro­tested, but the General Accounting Office decided on June 24, 1974, that “no reasonable basis" existed to question the validity of NASA’s decision. Thiokol, meanwhile, proceeded with design and develop­ment based on interim contracts, the final one for design awarded on June 26, 1974, followed by one for development, testing, and produc­tion on May 15, 1975.20

Part of the legacy from which Thiokol developed the technol­ogy for its SRBs came from the air force’s Large Segmented Solid Rocket Motor Program (designated 623A), of which Aerojet’s test­ing of 100-inch-diameter solids in the early 1960s had been an early part. In late 1962 the Air Force Rocket Propulsion Laboratory at Edwards AFB inaugurated a successor program. Its purpose was

to develop large solid motors that the DoD and NASA could use for space-launch vehicles. The air force provided funding for 120- and 156-inch-diameter segmented motors and for continuation of work on thrust-vector-control systems. NASA then paid for part of the 156-inch and all of a 260-inch program. In the course of test­ing thrust-vector-control systems, Lockheed had developed a Lock – seal mounting structure that allowed the nozzle to gimbal, and Thiokol later scaled it up to the size required for large motors, call – 270 ing it Flexseal.

Chapter 7 Lockheed tested both 120- and 156-inch motors in the program, and Thiokol tested 156-inch motors with both gimballed (Flexseal) and fixed nozzles. These tests concluded in 1967, as did those for 260-inch-diameter motors by Aerojet and Thiokol. There were no direct applications of the 260-inch technologies, but participation in the 120- and 156-inch portions of the Large Segmented Solid Rocket Motor Program gave Thiokol experience and access to designs, ma­terials, fabrication methods, and test results that contributed to de­velopment of the solid-rocket boosters for the Space Shuttle. The firm also drew upon its experience with Minuteman.21

The design for the solid-rocket booster was intentionally conser­vative, using a steel case of the same type (D6AC) used on Minute – man and the Titan IIIC. The Ladish Company of Cudahy, Wiscon­sin, made the cases for each segment without welding, using the rolled-ring forging process that it had helped develop for the Titan IIIC. In this process, technicians punched a hole in a hot piece of metal and then rolled it to the correct diameter. For the shuttle, the diameter turned out to be 12.17 feet (146 inches), with the overall length of the booster being 149 feet. Each booster consisted of four segments plus fore and aft sections. The propellant consisted of the same three principal ingredients used in the first stage of the Min – uteman missile, ammonium perchlorate, aluminum, and PBAN polymer. Its grain configuration was an 11-point star in the forward end converging into a large, smooth, tapered cylindrical shape. This combination yielded a theoretical specific impulse of more than 260 lbf-sec/lbm.22

Marshall Space Flight Center sought “to avoid inventing any­thing new" in the booster’s design, according to George Hardy, proj­ect manager for the solid-rocket booster at Marshall from 1974 to 1982. The best example of this approach was the PBAN propellant. Other propellants offered higher performance, but with cost and hu­man-rating being prime considerations, Thiokol employed a tried – and-true propellant used on the first stage of Minuteman and in the navy’s Poseidon missile. As Thiokol deputy director for the booster,

FIG. 7.4 Technical drawing of the Space Shuttle solid-rocket booster showing its segments and internal-burning core with other components, including its nozzle with gimbal actuators for directional (vector) control of the thrust. (Photo courtesy of NASA)

John Thirkill, said in 1973, “Over the last fifteen years, we’ve loaded more than 2,500 first stage Minuteman motors and around 500 Poseidon motors with this propellant."23

The configuration of the propellant grain caused the thrust to vary, providing the boost required for the planned trajectory but keeping the acceleration to 3 Gs for the astronauts. For the first six shuttle missions, the initial thrust was 3.15 million pounds per booster. The 11-point star in the forward section of the SRB had long, narrow points, providing an extensive burning surface. As the points burned away, the surface declined, reducing the thrust as the point of maximum dynamic pressure approached at about 60 sec­onds into the launch. At 52 seconds after liftoff, the star points had burned away to provide a cylindrical perforation in both the forward and rear segments of the booster. As this burned, expanding its di­ameter, the thrust increased slightly from the 52nd to about the 80th second. Thereafter, it tapered to zero as the burning consumed the propellant at about the 120th second, when the SRBs separated from the rest of the shuttle. The separated boosters, slowed by para­chutes, soon fell into the ocean.24

A major drawback of the PBAN propellant was that about 20 per­cent of its exhaust’s weight consisted of hydrogen chloride, which not

only was toxic and corrosive but could damage the ozone layer that protected Earth from excessive ultraviolet radiation. NASA studies of the possible ozone depletion showed, however, that it would be slight, so there was no need to shift to a less powerful propellant.25

Once the Ladish Company had forged the motor cases in Wiscon­sin, the segments traveled by railroad to a firm named Cal Doran near Los Angeles. There, heat treatment imparted greater strength and toughness to the D6AC steel. Then the segments went further 272 south to Rohr Industries in Chula Vista, near San Diego, for the Chapter 7 addition of tang-and-clevis joints to the ends of the segments. On these joints, shuttle designers had departed from the Marshall ad­vice “to avoid inventing anything new." Although the shuttle field joints resembled those for Titan IIIC, in many respects they dif­fered. One key change lay in orientation. For the Titan solid-rocket motor, the single tang pointed upward from a lower segment of the case and fit into the two-pronged clevis, which encased it. This pro­tected the joint from rain or dew dripping down the case and enter­ing the joint. In the shuttle, the direction was reversed.

A second major difference lay in the Titan joint’s having used only one O-ring, whereas the shuttle employed two. Insulation on the inside of the Titan motor case protected the case, and with it, the O-ring, from excessive heating. To keep the protective mecha­nisms from shrinking in cold temperatures and then possibly al­lowing a gas blow-by when the motor was firing, there were heating strips on the Titan. Both the Titan and the shuttle used putty to improve the seal provided by the O-ring(s), but the shuttle added the second O-ring for supposed further insurance. It did not include heating strips, however. One further difference in the joints was in the number of pins holding the tang and clevis together. Whereas the Titan motor had used 240 such pins fitting into holes in the tang and clevis and linking them, the shuttle had only 177, despite its larger diameter.26 There is no certainty in counterfactual history, but perhaps if the shuttle designers had simply accepted the basic design of the Titan tang-and-clevis joints, the Challenger accident would not have occurred because of leaking hot gases through a field joint that ignited the external tank.

Unlike the field joint, the nozzle for the solid-rocket boosters did follow the precedents of the Titan solid-rocket motors and the Large Segmented Solid Rocket Motor Program. The shuttle employed car­bon-phenolic throats to ablate under the extreme heating from the flow and expansion of the hot gases from the burning propellant in the motor itself. In the case of the shuttle, the propellants burned at a temperature of 5,700°F, so ablation was needed to vaporize and

thereby prevent thermal-stress cracking followed by probable ejec­tion of portions of the nozzle. As of June 1979, the expansion ratio of the nozzle was 7.16:1, used for the first seven missions. Start­ing with the eighth mission, modifications of the nozzle increased the initial thrust of each motor from 3.15 million to 3.3 million pounds. These changes extended the length of the nozzle exit cone by 10 inches and decreased the diameter of the nozzle throat by 4 inches. The latter change increased the expansion ratio to 7.72:1, thereby adding to the booster’s thrust.27

The nozzle was partially submerged, and for gimballing, it used the Flexseal design Thiokol had scaled up in the 156-inch motor testing from the Lockheed’s Lockseal design. It was capable of eight degrees of deflection, necessitated among other reasons by the shuttle’s now-familiar roll soon after liftoff to achieve its proper trajectory. Having less thrust, the space shuttle main engines were incapable of achieving the necessary amount of roll, and the liq­uid-injection thrust-vector-control system used on the Titan solid – rocket motors would not have met the more demanding require­ments of the shuttle. Hence the importance of the Lockseal-Flexseal development during the Large Segmented Solid Rocket Motor Pro­gram supported by both NASA and the air force.28

Although there were only four segments of the solid-rocket boosters that were joined by field joints, there were actually 11 sec­tions joined by tang-and-clevis joints. Once they had been through machining and fitting processes, they were assembled at the factory into four segments. The joints put together at the factory were called factory joints as distinguished from the field joints, which techni­cians assembled at Kennedy Space Center. Thiokol poured and cast the propellant into the four segments at its factory in Brigham City, Utah, usually doing so in matched pairs from the same batches of propellant to reduce thrust imbalances. At various times, the solid – rocket motors used four different D6AC-steel cases, with slight variations in thickness.29

In part because of its simplicity compared with the space shut­tle main engine, the solid-rocket booster required far less testing than the liquid-propellant engine. Certification for the SSMEs had required 726 hot-fire tests and 110,000 seconds of operation, but the solid-rocket boosters needed only four developmental and three qualification tests with operation of less than 1,000 seconds total— 0.9 percent of that for the SSMEs. There were, however, other tests. One was a hydroburst test on September 30, 1977, at Thiokol’s Wa­satch Division in Utah. This demonstrated that, without cracking, a case could withstand the pressures to which it would be subjected

during launch. A second hydroburst test on Sept ember 19, 1980 (with only the aft dome, two segments, and the forward dome), was also successful. There were other tests of the tang-and-clevis joints that put them under pressure until they burst. They withstood pres­sures between 1.72 and 2.27 times the maximum expected from liftoff through separation.30

The first developmental static test, DM-1 on July 18, 1977, at Thiokol’s Wasatch Division was successful, but the motor deliv – 274 ered only 2.9 million pounds of maximum thrust compared with Chapter 7 an expected 3.1 million. There were other anomalies, including ex­cessive erosion in parts of the nozzle. Modification included addi­tional ammonium perchlorate in the propellant and changed nozzle coatings. DM-2 on January 18, 1978, was another success but led to further adjustments in the design. It turned out that the rubber insulation and polymer liner protecting the case were thicker than necessary, leading to reduction in their thickness. This lowered their weight from 23,900 to 19,000 pounds. There were also modi­fications in the igniter, grain design, and nozzle coating to reduce the flame intensity of the igniter, the rate of thrust increase for the motor, and erosion of portions of the nozzle. As the motor for DM-3 was being assembled, a study of the DM-2 casing revealed that there had been an area with propellant burning between segments. This required disassembling the motor and increasing the thickness of a noncombustible inhibitor on the end of each segment. Designers also extended the rubber insulation to protect the case at the joints. This delayed the DM-3 test from July to October 19, 1978.

Again, the test was satisfactory; but although the thermal protec­tion on the nozzle had been effective, the igniter once more caused the thrust to rise too quickly. Designers could see no evident solu­tion to the rapid rate of thrust increase, an apparent tacit admis­sion that engineers did not fully understand the complex combus­tion process. It did seem evident, though, that the rate had to rise quickly to preclude thrust imbalances between the two motors, so the engineers went back to an igniter design closer to that used in the DM-1 test and simply accepted the rapid thrust rise (for the moment, at least). On February 17, 1979, DM-4 ended the four de­velopmental tests with a successful firing. The qualification tests, QM-1 through -3 from June 13, 1979, to February 13, 1980, were all successful. These seven tests furnished the data needed to qualify the solid-rocket motor for launch—excluding the electronics, hy­draulics, and other components not Thiokol’s responsibility. Other tests on booster recovery mechanisms, complete booster assem­blies, loads on the launch pad and in flight, and internal pressure

FIG. 7.5 Testing of a developmental motor following the Challenger accident. (Photo courtesy of NASA)

took place at Marshall and at the National Parachute Test Range, El Centro, California. The program completed all of these tests by late May 1980, well before the first shuttle flight.31 Of course, this was after the first planned flight, so if the main-engine development had not delayed the flights, presumably the booster development would have done so to some degree.

Smaller Solid-Propellant Stages and Boosters

Even early in launch-vehicle history, some missile programs had al­ready begun to influence solid-propellant developments. In 1956, a creative group of engineers at Langley’s Pilotless Aircraft Research Division (PARD) began formulating ideas that led to the Scout launch vehicles. This group included Maxime A. Faget, later famous for designing spacecraft; Joseph G. Thibodaux Jr., who promoted the spherical design of some rocket and spacecraft motors beginning in 1955; Robert O. Piland, who put together the first multistage rocket to reach the speed of Mach 10; and William E. Stoney Jr., who be­came the first head of the group responsible for developing the Scout, which he also christened. Wallops, established as a test base for the National Advisory Committee for Aeronautics’ (NACA) Langley Memorial Aeronautical Laboratory in 1945, had a history of using rockets, individually or in stages, to gather data at high speeds on both aircraft models and rocket nose cones. These data made it pos­sible to design supersonic aircraft and hypersonic missiles at a time when ground facilities were not yet capable of providing comparable information. It was a natural step for engineers working in such a pro­gram to conceive a multistage, hypersonic, solid-propellant rocket that could reach orbital speeds of Mach 18.32