Propulsion Engineering Innovations Thermal Protection Systems
Materials and Manufacturing
Aerodynamics and Flight Dynamics
Avionics, Navigation, and Instrumentation
Software
Structural Design
Robotics and Automation
Systems Engineering for Life Cycle of Complex Systems
Engineering Innovations 157 Propulsion The launch of the Space Shuttle was probably the most visible event of the entire mission cycle. The image of the Main Propulsion System— the Space Shuttle Main Engine and the Solid Rocket Boosters (SRBs)— Introduction powering the Orbiter into space captured the attention and the Yolanda Harris imagination of people around the globe. Even by 2010 standards, Space Shuttle Main Engine these main engine s’ performance was unsurpassed compared to any Fred Jue other engines. They were a quantum leap from previous rocket engines. Chemochromic Hydrogen Leak Detectors The main engines were the most reliable and extensively tested rocket Luke Roberson engine before and during the shuttle era. Janine Captain Martha Williams The shuttl e’s SRBs were the largest ever used, the first reusable rocket, Mary Whitten and the only solid fuel certified for human spaceflight. This technology, The First Human-Rated engineering, and manufacturing may remain unsurpassed for decades Reuseable Solid Rocket Motor Fred Perkins to come. Holly Lamb But the shuttl e’s propulsion capabilities also encompassed the Orbite r’s Orbital Propulsion Systems Cecil Gibson equally important array of rockets—the Orbital Maneuvering System Willard Castner and the Reaction Control System—which were used to fine-tune orbits Robert Cort and perform the delicate adjustments needed to dock the Orbiter Samuel Jones with the International Space Station. The design and maintenance of Pioneering Inspection Tool the first reusable space vehicle—the Orbiter—presented a unique set Mike Lingbloom of challenges. In fact, the Space Shuttle Program developed the worl d’s Propulsion Systems and Hazardous Gas Detection most extensive materials database for propulsion. In all, the shuttl e’s Bill Helms propulsion systems achieved unprecedented engineering milestones and David Collins launched a 30-year era of American space exploration. Ozzie Fish Richard Mizell
158 Engineering Innovations Space Shuttle
Main Engine S Spa ce Shu ttle M ain En gine Prop ellan t Flow
NASA faced a unique challenge at Hydrydrogenogen Oxygen the beginning of the Space Shuttle Inlet Inlet LLow-pressureow-pressure LLow-pressureow-preessurssure Program: to design and fly a Fuel TurbopumpTurbopump Oxidizer TurbopumpTurbopump human-rated reusable liquid propulsion Main Oxidizer Oxidizer rocket engine to launch the shuttle. VValvealve ValveValvalve It was the first and only liquid-fueled rocket engine to be reused from Oxidizer VValvealve Oxidizer one mission to the next during the PPreburnerreburner shuttle era. The improvement of the Fuel Main Space Shuttle Main Engine (SSME) PreburnerPreburner Injector was a continuous undertaking, Powerhead with the objectives being to increase safety, reliability, and operational Main margins; reduce maintenance; and Fuel ValveValve Main improve the life of the engine’s Combustion High-prigh-pressureessure Chamber Oxidizer TTurbopumpurbopump high-pressure turbopumps. Nozzle HHigh-pressureigh-pressure FFueluel TTurbopumpurbopump The reusable SSME was a staged . d e v r combustion cycle engine. Using a e s
Chamber e r
s mixture of liquid oxygen and liquid Coolant VValvealvalve t h g i r
l hydrogen, the main engine could attain l A
. e n
a maximum thrust level (in vacuum) y d t e of 232,375 kg (512,300 pounds), k c o R The Space Shuttle Main Engine used a two-stage combustion process. Liquid hydrogen which is equivalent to greater than y e n t and liquid oxygen were pumped from the External Tank and burned in two preburners. i 12,000,000 horsepower (hp). The h W The hot gases from the preburners drove two high-pressure turbopumps—one for liquid &
t engine also featured high-performance t a
hydrogen (fuel) and one for liquid oxygen (oxidizer). r P fuel and oxidizer turbopumps that © developed 69,000 hp and 25,000 hp, respectively. Ultra-high-pressure by the new Space Transportation The design team chose a dual-preburner operation of the pumps and combustion System (STS). powerhead configuration to provide chamber allowed expansion of hot precise mixture ratio and throttling In 1971 , the Rocketdyne division of gases through the exhaust nozzle to control. A low- and high-pressure Rockwell International was awarded a achieve efficiencies never previously turbopump, placed in series for each of contract to design, develop, and attained in a rocket engine. the liquid hydrogen and liquid oxygen produce the main engine. loops, generated high pressures across a Requirements established for Space The main engine would be the first wide range of power levels. Shuttle design and development began production-staged combustion in the mid 1960s. These requirements A weight target of 2,857 kg (6,300 cycle engine for the United States. called for a two-stage-to-orbit vehicle pounds) and tight Orbiter ascent (The Soviet Union had previously configuration with liquid oxygen envelope requirements yielded a demonstrated the viability of staged (oxidizer) and liquid hydrogen (fuel) compact design capable of generating combustion cycle in the Proton vehicle for the Orbiter’s main engines. By a nominal chamber pressure of in 1965.) The staged combustion 1969, NASA awarded advanced engine 211 kg/cm 2 (3,000 pounds/in 2)—about cycle yielded high efficiency in a studies to three contractor firms to four times that of the Apollo/Saturn technologically advanced and complex further define designs necessary to J-2 engine. engine that operated at pressures meet the leap in performance demanded beyond known experience.
Engineering Innovations 159 For the first time in a boost-to-orbit rocket engine application, an on-board digital main engine controller continuously monitored and controlled all engine functions. The controller Michael Coats Pilot on STS-41D (1984). initiated and monitored engine Commander on STS-29 (1989) parameters and adjusted control and STS-39 (1991). valves to maintain the performance parameters required by the mission. When detecting a malfunction, it also commanded the engine into a safe A Balky Hydrogen Valve lockup mode or engine shutdown. Halts Discovery Liftoff
“I had the privilege of being the pilot on the maiden flight of the Orbiter Design Challenges Discovery, a hugely successful mission. We deployed three large communications Emphasis on fatigue capability, satellites and tested the dynamic response characteristics of an extendable strength, ease of assembly and solar array wing, which was a precursor to the much-larger solar array wings disassembly, maintainability, and materials compatibility were all major on the International Space Station. considerations in achieving a fully “But the first launch attempt did not go quite as we expected. Our pulses were reusable design. racing as the three main engines sequentially began to roar to life, but as we Specialized materials needed to be rocked forward on the launch pad it suddenly got deathly quiet and all motion incorporated into the design to meet the severe operating environments. NASA stopped abruptly. With the seagulls screaming in protest outside our windows, successfully adapted advanced alloys, it dawned on us we weren’t going into space that day. The first comment including cast titanium, Inconel ® 718 came from Mission Specialist Steve Hawley, who broke the stunned silence (a high-strength, nickel-based superalloy by calmly saying ‘I thought we’d be a lot higher at MECO (main engine cutoff).’ used in the main combustion chamber So we soon started cracking lousy jokes while waiting for the ground crew support jacket and powerhead), and to return to the pad and open the hatch. The joking was short-lived when NARloy-Z (a high-conductivity, copper-based alloy used as the liner in we realized there was a residual fire coming up the left side of the Orbiter, fed the main combustion chamber). NASA from the same balky hydrogen valve that had caused the abort. The Launch also oversaw the development of Control Center team was quick to identify the problem and initiated the water single-crystal turbine blades for the deluge system designed for just such a contingency. We had to exit the pad high-pressure turbopumps. This elevator through a virtual wall of water . We wore thin, blue cotton flight suits innovation essentially eliminated the grain boundary separation failure back then and were soaked to the bone as we entered the air-conditioned mechanism (blade cracking) that had astronaut van for the ride back to crew quarters. Our drenched crew shivered limited the service life of the pumps. and huddled together as we watched the Discovery recede through the rear Nonmetallic materials such as Kel-F ® window of the van, and as Mike Mullane wryly observed, ‘This isn’t exactly (a plastic used in turbopump seals), ® what I expected spaceflight to be like.’ The entire crew, including Commander Armalon fabric (turbopump bearing cage material), and P5N carbon-graphite Henry Hartsfield, the other Mission Specialists Mike Mullane and Judy Resnik, seal material were also incorporated and Payload Specialist Charlie Walker, contributed to an easy camaraderie that into the design. made the long hours of training for the mission truly enjoyable.” Material sensitivity to oxygen environment was a major concern for compatibility due to reaction and
160 Engineering Innovations ignition under the high pressures. Mechanical impact testing had vastly expanded in the 1970s to accommodate the shuttle engine’s varied operating conditions. This led to a new class of liquid oxygen reaction testing up to 703 kg/cm 2 (1 0,000 pounds/in 2). . d e v r e
Engineers also needed to understand s e r
s t
long-term reaction to hydrogen effects h g i r
l l
to achieve full reusability. Thus, a A
. e n whole field of materials testing evolved y d t e k to evaluate the behavior of hydrogen c o R
y
charging on all affected materials. e n t i h W
NASA developed new tools to &
t t a r
accomplish design advancements. P
Engineering design tools advanced © along with the digital age as analysis A 1970s-era Space Shuttle Main Engine undergoes testing at Rocketdyne’s Santa Susana Field migrated from the mainframe platform Laboratory near Los Angeles, California. to workstations and desktop personal computers. Fracture mechanics and Test Bed—occurred in 1975 at the related to the initial design of the fracture control became critical tools NASA National Space Technology high-pressure turbopumps, powerhead, for understanding the characteristics of Laboratory (now Stennis Space Center) valves, and nozzles. in Mississippi and relied on facility crack propagation to ensure design Extensive margin testing beyond the controls, as the main engine controller reusability. As the analytical tools and normal flight envelope—including was not yet available. processor power improved over the high-power, extended-duration tests and decades, cycle time for engineering NASA and Rocketdyne pursued an near-depleted inlet propellant conditions analysis such as finite element models, aggressive test schedule at their to simulate the effects of microgravity— computer-aided design and respective facilities. Stennis Space provided further confidence in the manufacturing, and computational fluid Center with three test stands and design. Engineers subjected key dynamics dropped from days to minutes. Rocketdyne with one test stand components to a full series of design Real-time engine performance analyses completed 152 engine tests in 1980 verification tests, some with intentional were conducted during ground tests and alone—a record that has not been hardware defects, to validate safety flights at the end of the shuttle era. exceeded since. This ramp-up to margins should the components develop 100,000 seconds represented a team undetected flaws during operation. effort of personnel and facilities to Development and Certification NASA and Rocketdyne also overachieve a stated development performed system testing to replicate The shuttle propulsion system was goal of 65,000 seconds set by the three engine cluster interactions the most critical system during then-Administrator John Yardley as with the Orbiter. The Main Propulsion ascent; therefore, a high level of the maturity level deemed flightworthy. Test Article consisted of an Orbiter testing was needed prior to first flight NASA verified operation at altitude aft fuselage, complete with full thrust to demonstrate engine maturity. conditions and also demonstrated the structure, main propulsion electrical Component-level testing of the rigors of sea-level performance and and system plumbing, External Tank, preburners and thrust chamber began engine gimballing for thrust vector and three main engines. To validate that in 1974 at Rocketdyne’s Santa Susana control. The Rocketdyne laboratory the Main Propulsion System was ready Field Laboratory in Southern California. supplemented sea-level testing as well for launch, engineers completed 18 tests as deep throttling by using a low 33:1 The first engine-level test of the main at the National Space Technology expansion ratio nozzle. This testing was engine—the Integrated Subsystem Laboratory by 1981 . crucial in identifying shortcomings
Engineering Innovations 161 The completion of the main engine routine engine maintenance without chamber, and high-pressure oxidizer preliminary flight certification in removing them from the Orbiter. and fuel turbopumps. March 1981 marked a major milestone The successful flight of STS- 1 initiated These major changes would later be in clearing the initial flights at 100% the development of a full-power (1 09% divided into two “Block” configuration rated power level. rated power level) engine. The higher upgrades, with Rocketdyne tasked to thrust capability was needed to support improve the powerhead, heat exchanger, Design Evolutions an envisioned multitude of NASA, and main combustion chamber while commercial, and Department of Defense Pratt & Whitney was selected to design, A major requirement in engine design payloads, especially if the shuttle was develop, and produce the improved was the ability to operate at various launched from the West Coast. By 1983, high-pressure turbopumps. power levels. The original engine life however, test failures demonstrated the Pratt & Whitney Company of United requirement was 100 nominal missions basic engine lacked margin to Technologies began the effort in 1986 and 27,000 seconds (7.5 hours) of continuously operate at 109% thrust, and to provide alternate high-pressure engine life. Nominal thrust, designated full-power-level development was turbopumps as direct line replaceable as rated power level, was 21 3, 189 kg halted. Other engine improvements were units for the main engines. Pratt & (470,000 pounds) in vacuum. The life implemented into what was called the Whitney used staged combustion requirement included six exposures Phase II engine. During this period, the experience from its development of the at the emergency power level of engine program was restructured into XL R-129 engine for the US Air Force 232,375 kg ( 51 2,300 pounds), which two programs—flight and development. was designated 109% of rated power and cryogenic hydrogen experience level. To maximize the number of Post-Challenger Return to Flight from the RL -1 0 (an upper- stage engine missions possible at emergency power used by NASA , the military, and level, an assessment of the engine The 1986 Challenger accident provoked commercial enterprises) along with capability resulted in reducing the fundamental changes to the shuttle, SSME lessons learned to design the number of nominal missions per engine including an improved main engine new pumps. The redesign of the to 55 missions at 109%. Emergency called Phase II. This included changes components eliminated critical failure power level was subsequently renamed to the high-pressure turbopumps and modes and increased safety margins. full power level. main combustion chamber, avionics, valves, and high-pressure fuel duct Next Generation Ongoing ascent trajectory analysis insulation. An additional 90,2 41 The Block I configuration became determined 65% of rated power level seconds of engine testing accrued, the successor to the Phase II engine. to be sufficient to power the vehicle including recertification to 104% rated A new Pratt & Whitney high-pressure through its period of maximum power level. aerodynamic pressure during ascent. oxygen turbopump, an improved Minimum power level was later refined The new Phase II engine continued two-duct engine powerhead, and upward to 67%. to be the workhorse configuration a single-tube heat exchanger were for shuttle launches up to the late introduced that collectively used On April 12, 1981 , Space Shuttle 1990s while additional improvements new design and production processes Columbia lifted off Launch Pad 39A envisioned during the 1980s were to eliminate failure causes. Also it from Kennedy Space Center in Florida undergoing development and flight increased the inherent reliability on its maiden voyage. The first flight certification for later incorporation. and operating margin and reduced configuration engines were aptly named NASA targeted five major components production cycle time and costs. the First Manned Orbital Flight SSMEs. for advanced development to further This Block I engine first flew on These engines were flown during the enhance safety and reliability, STS-70 (1 995). initial five shuttle development missions lower recurring costs, and increase The powerhead redesign was less at 100% rated power level thrust. performance capability. These risky and was chosen to proceed ahead Work done to prepare for the next components included the powerhead, of the main combustion chamber. flight validated the ability to perform heat exchanger, main combustion
162 Engineering Innovations As Block II development testing progressed, the engineering accomplishments on the large-throat main combustion chamber matured more rapidly than the high-pressure fuel turbopump. By February 1997, NASA had decided to go forward with an interim configuration called the Block IIA. . d e
v Using the existing Phase II r e s e r
high-pressure fuel pump, this s t h g i
r configuration would allow early
l l A
. implementation of the large-throat e n y d t main combustion chamber to support e k c o ISS launches. The large-throat main R
y e n
t combustion chamber was simpler i h W and producible. The new chamber &
t t a r lowered the engine’s operating P
© pressures and temperatures while The Technology Test Bed Space Shuttle Main Engine test program was conducted at Marshall Space increasing the engine’s operational Flight Center, Alabama, between September 1988 and May 1996. The program demonstrated the ability safety margin. Changes to the of the main engine to accommodate a wide variation in safe operating ranges. low-pressure turbopumps to operate in this derated environment, along The two-duct powerhead eliminated component-level testing occurred at the with further avionics improvements, 74 welds and had 52 fewer parts. Pratt & Whitney West Palm Beach, were flown in 1998 on STS-89. This improved design led to production Florida, testing facilities. Testing then simplification and a 40% cost reduction graduated to the engine level at Stennis The large-throat main combustion compared to the previous three-duct Space Center as well as at Marshall chamber became one of the most configuration. The two-duct Space Flight Center’s (MSFC’s) significant safety improvements for configuration provided an improvement Technology Test Bed test configuration. the main engine by effectively reducing to the hot gas flow field distribution and operating pressures and temperatures The large-throat main combustion reductions in dynamic pressures. The up to 10% for all subsystems. This chamber began prototype testing at improved heat exchanger eliminated design also incorporated improved Rocketdyne in 1988. But it was all inter-propellant welds, and its wall cooling capability for longer life and not until 1992, after a series of thickness was increased by 25% for used high-strength castings, thus combustion stability tests at the added margin against penetration by eliminating 50 welds. MSFC Technology Test Bed facility, unexpected foreign debris impact. that concerns regarding combustion By the time the first Block IIA flew on The new high-pressure oxygen stability were put to rest. The next STS-89 in January 1998, the large-throat turbopump eliminated 293 welds, added improved engine—Block II— main combustion chamber design had improved suction performance, and incorporated the new high-pressure accumulated in excess of 100,000 introduced a stiff single-piece disk/shaft fuel turbopump, modified low-pressure seconds of testing time. By late 1999 , the configuration and thin-cast turbine turbopumps, software operability Block II high-pressure fuel turbopump blades. The oxygen turbopump enhancements, and previous Block I had progressed into certification testing. incorporated silicon nitride (ceramic) upgrades. These upgrades were The design philosophy mirrored ball bearings in a rocket engine needed to support International Space those proven successful in the application and could be serviced Station (ISS) launches with their heavy high-pressure oxidizer turbopump and without removal from the engine. Initial payloads beginning in 1998. included the elimination of 387 welds
Engineering Innovations 163
The Improved Space Shuttle Main Engine Powerhead Component Arrangement for Block II Engines
Fuel Oxidizer Preburnerreburner Preburnerreburner
MainMain Combustion Chamber High-pressureHigh-pressure High-pressureHHigh-prigh-pressure Fuel TurbopumpTurbopump OxidizerOxidizer TurbopumpTurbopump
The Block II engine combined a new high-pressure fuel turbopump with the previously flown redesigned high-pressure oxygen turbopump. DRAFT 8/20/09 Risk analysis showed that the Block II engine was twice as safe as the 1990s-era engine. Beginning with STS-110 in April 2002, all shuttle flights were powered by the improved Space Shuttle Main Engine.
and incorporation of a stiff single-piece projected that the Block II engine was the system further improved engine disk/shaft, thin-cast turbine blades, twice as safe as the Phase II engine. ascent safety by an additional 23%. and a cast pump inlet that improved the The first two single-engine flights of suction performance and robustness Block II occurred on STS -1 04 and Summary against pressure surges. As with the STS -1 08 in July 2001 and December high-pressure oxidizer turbopump, Another major SSME milestone took 20 01 , respectively, followed by the first the high-pressure fuel turbopump place in 2004 when the main engine three-engine cluster flight on STS -11 0 turbine inlet did not require off-engine passed 1,000,000 seconds in test and in April 2002. The high-pressure fuel inspections, which contributed operating time. This unprecedented turbopump had accumulated 150,843 significantly to improving engine level of engine maturity over the seconds of engine test maturity at the turnaround time. The high-pressure preceding 3 decades established the time of the first flight. fuel turbopump also demonstrated main engine as one of the world’s that a turbine blade failure would result The Block II engine also incorporated most reliable rocket engines, with a in a contained, safe engine shutdown. the advanced health management 100% flight safety record and a By introducing the added operational system on STS -11 7 in 2007. This demonstrated reliability exceeding margin of the large-throat main on-board system could detect and 0.9996 in over 1,000,000 seconds of combustion chamber with the new mitigate anomalous high-pressure hot-fire experience. turbopumps, quantitative risk analysis turbopump vibration behavior, and
164 Engineering Innovations The First Human- Chemochromic Hydrogen Leak Detectors Rated Reusable Solid Rocket Motor The Chemochromic Point Detector for sensing hydrogen gas leakage is useful in any application in which it is important to know the presence and location of a The Space Shuttle reusable solid hydrogen gas leak. rocket motors were the largest solid rockets ever used, the first reusable This technology uses a chemochromic pigment and polymer that can be molded or spun solid rockets, and the only solids ever into a rigid or pliable shape useable in variable-temperature environments including certified for crewed spaceflight. The atmospheres of inert gas, hydrogen gas, or mixtures of gases. A change in the color of closest solid-fueled rival—the Titan IV detector material reveals the location of a leak. Benefits of this technology include: Solid Rocket Motor Upgrade—was known for boosting heavy payloads temperature stability, from -75°C to 100°C (-103°F to 212°F); use in cryogenic for the US Air Force and National applications; ease of application and removal; lack of a power requirement; quick Reconnaissance Organization. The response time; visual or electronic leak detection; nonhazardous qualities, thus motors were additionally known for environmentally friendly; remote monitoring capability; and a long shelf life. This launching the 5,586-kg (1 2,220-pound) technology is also durable and inexpensive. Cassini mission on its 7-year voyage to Saturn. By contrast, the Titan The detector can be fabricated into two types of sensors—reversible and irreversible. booster was 76 cm (30 in.) smaller in Both versions immediately notify the operator of the presence of low levels of diameter and 4.2 m (1 4 ft) shorter in length, and held only two-thirds of the hydrogen; however, the reversible version does not require replacement after exposure. amount of propellant. Both versions were incorporated into numerous polymeric materials for specific In a class of its own, the Reusable applications including: extruded tapes for wrapping around valves and joints suspected Solid Rocket Motor Program was of leaking; injection-molded parts for seals, O-rings, pipe fittings, or plastic piping characterized from its inception by four material; melt-spun fibers for clothing applications; and paint for direct application distinguishing traits: hardware to ground support equipment. The versatility of the sensor for several different reusability, postflight recovery and applications provides the operator with a specific-use safety notification while working analysis, a robust ground-test program, and a culture of continual improvement under hazardous operations. via process control. The challenge NASA faced in developing the first human-rated solid rocket motor was to engineer a pair of solid-fueled rocket motors capable of meeting the rigorous reliability requirements associated with human spaceflight. The rocket motors would Hydrogen-sensing tape applied to the have to be powerful enough to boost the Orbiter midbody umbilical unit during fuel shuttle system into orbit. The motors cell loading for STS-118 through STS-123 would also need to be robust enough to at Kennedy Space Center, Florida. meet stringent reliability requirements and survive the additional rigors of Hydrogen-sensing tape application at re-entry into Earth’s atmosphere and liquid hydrogen cross-country vent line subsequent splashdown, all while being flanges on the pad slope. reusable. The prime contractor— Morton Thiokol, Utah—completed its
Engineering Innovations 165 the steel could be heated and melted by the 2,760°C (5,000°F) combustion gases. Too much insulation, and weight requirements were exceeded. Engineers employed sophisticated design analysis and testing to optimize this balance between protection and weight. By design, much of the insulation was burned away during the 2 minutes of motor operation. The propellant was formulated from . d
e three major ingredients: aluminum v r e s
e powder (fuel); ammonium perchlorate r
s t
h (oxidizer); and a synthetic polymer g i r
l l A
binding agent. The ingredients were . K T A
batched, fed into large 2,600-L © (600-gal) mix bowls, mixed, and tested The two shuttle reusable solid rocket motors, which stood more than 38 m (126 ft) tall, harnessed before being poured into the insulated 29.4 meganewtons (6.6 million pounds) of thrust. The twin solid-fueled rockets provided 80% of the thrust needed to achieve liftoff. and lined segments. Forty batches were produced to fill each case first full-scale demonstration test A Proven Design segment. The propellant mixture had within 3 years. an initial consistency similar to that of To construct the reusable solid rocket peanut butter, but was cured to a texture NASA learned a poignant lesson in motor, four cylindrical steel segments— and color that resembled a rubber the value of spent booster recovery insulated and loaded with a high- pencil eraser—strong, yet pliable. and inspection with the Challenger performance solid propellant—were The propellant configuration or “shape” tragedy in January 1986. The joined together to form what was inside each segment was carefully postflight condition of the hardware essentially a huge pressure vessel and designed and cast to yield the precise provided valuable information on the combustion chamber. The segmented thrust trace upon ignition. health of the design and triggered a design provided maximum flexibility in Once each segment was insulated redesign effort that surpassed, in motor fabrication, transportation, and and cast with propellant and finalized, magnitude and complexity, the original handling. Each segment measured 3.7 m the segments were shipped from development program. (1 2 ft) in diameter and was forged from ATK’s manufacturing facility in D6AC steel measuring approximately For the substantial redesign that Utah to Kennedy Space Center (KSC) 1.27 cm (0.5 in.) in thickness. occurred between 1986 and 1988, in Florida, on specially designed, engineers incorporated lessons learned Case integrity and strength were heavy-duty covered rail cars. At KSC, from the first 25 shuttle flight booster maintained during flight by insulating they were stacked and assembled into sets. More than 100 tests, including the case interior. The insulating liner was the flight configuration. five full-scale ground tests, were a fiber-filled elastomeric (rubber-like) The segments were joined together conducted to demonstrate the strength material applied to the interior of the with tang/clevis joints pinned in 177 of the new design. Flaws were steel cylinders. A carefully formulated locations and sealed with redundant deliberately manufactured into the final tacky rubber bonding layer—or O-rings. Each joint, with its redundant test motor to check redundant systems. “liner”—was applied to the rubber seals and multiple redundant seal insulator surface to facilitate a strong The redesigned motors flew for the protection features, was pressure bond with the propellant. first time in September 1988 and checked during assembly to ensure a performed flawlessly. Producing an accurate insulating layer good pressure seal. was critical. Too little insulation, and
166 Engineering Innovations An igniter was installed in the forward end of the forward segment— Reusable Solid Rocket Motor Propellant Configuration at the top of the rocket. The igniter was essentially a smaller rocket motor that fired into the solid rocket motor Forward to ignite the main propellant grain. Design and manufacture closely mirrored the four main segments. Forward Casting Center Segments The nozzle was installed at the aft end of the aft segment, at the bottom of the Aft Center rocket. The nozzle was the “working” component of the rocket in which hot Aft exhaust gases were accelerated and
Nozzle Protective Plug directed to achieve performance requirements and vehicle control.
Propellant The nozzle structure consisted of metal housings over which were bonded layers of carbon/phenolic and .
Nozzle d silica/phenolic materials that protected e v
Aft Exit Cone r e
s the metal structure from the searing e r
s t exhaust gases by partially decomposing h
The four primary propulsion segments that comprised the reusable solid rocket motor g i r
l l and ablating. A flexible bearing, formed were manufactured individually then assembled for launch. Each segment was reusable A
. K and designed for a service life of up to 20 flights. T with vulcanized rubber and steel, A
© allowed for nozzle maneuverability up to 8 degrees in any direction to steer the shuttle during the first minutes of flight. F Forward Segment Propellant Grain Configuration Engineers employed significant analysis and testing to develop a reliable and efficient nozzle capable Fin Mold Line Transition Region Castable of being manufactured. The nozzle Center-perforated Inhibitor Bore Region flexible bearing—measuring up Liner 2.35 m (92.4 in.) at its outside (hand applied) Star Region Propellant Liner (sling applied) diameter—was an example of one component that required multiple Dome Case Nitrate Butadiene processing iterations to ensure Rubber Insulation Fin Tip the manufactured product aligned Fin Cavity with design requirements. Star Point NASA enhanced the nozzle design Center-perforated Bore following the Challenger accident when severe erosion on one section of the nozzle on one motor was noted through . d
e postflight analysis. While the phenolic v r e s e
r liners were designed to erode smoothly
s t h g
i and predictably, engineers found—at r
l l A
The forward propulsion segment featured a unique grain pattern designed to yield the . certain ply orientations—that internal K greatest thrust when it was needed most—on ignition. T A stresses resulting from exposure to hot ©
Engineering Innovations 167 gases exceeded the material strength. over time, changes were inevitable. Postflight Analysis Under such stress, the hot charred Change to design or process became The ability to closely monitor flight material had the potential to erode mandatory as a result of factors such performance through hands-on erratically and jeopardize component as material/vendor obsolescence or new postflight analysis—after myriad integrity. Engineers modified nozzle environmental regulations. material, design, and process changes— ply angels to reduce material stress, was only possible by virtue of the and this condition was successfully Changing Processes motor’s reusable nature. eliminated on all subsequent flights. During a 10-year period beginning Developing methods to scrutinize in the mid 1990s, for example, more and recertify spent rocket motor than 100 supplier materials used to hardware that had raced through the produce the reusable solid rocket stratosphere at supersonic speeds was motor became obsolete. The largest new. NASA had the additional burden contributing factor stemmed from of working with components that had supplier economics, captured in three
. experienced splashdown loads and d
e main scenarios. First, suppliers v r
e were subsequently soaked in corrosive s
e changed their materials or processes. r
s
t saltwater prior to retrieval. h Second, suppliers consolidated g i r
l l A
operations and either discontinued or
. In the early days of the program, K T A otherwise modified their materials. NASA made significant efforts in © Technicians shown installing igniter used to Third, the materials were simply no identifying relevant evaluation criteria initiate the propellant burn in a forward motor longer available from subtier vendors. and establishing hardware assessment segment. The igniter was a small rocket methods. A failure to detect hardware US environmental regulations, such motor loaded with propellant that propagated stresses and material weaknesses could flame down the bore of the motor. as the requirement to phase out the result in an unforgivable catastrophic use of ozone-depleting chemicals, event later on. The criteria used to were an additional factor. Methyl evaluate the first motors and the The Reusable Rocket chloroform, for example, was a solvent accompanying data collected would All metal hardware—including used extensively in hardware also become the benchmark from structures from the case, igniter, processing. A multimillion-dollar which future flights would be measured. safe-and-arm device, and nozzle— effort was launched within NASA and Included in the evaluation criteria were designed to support up to 20 ATK to eventually eliminate methyl were signs of case damage or material shuttle missions. This was unique to chloroform use altogether in motor loss caused by external debris; integrity the reusable solid rocket motor. processing. Eight alternate materials of major components such as case Besides the benefits of conservation were selected following thorough segments, nozzle and igniter; and and affordability, the ability to testing and analysis to ensure program fidelity of insulation, seals, and joints. recover the motors allowed NASA performance was not compromised. to understand exactly how the Inspection and documentation of components performed in flight. New Technology retrieved hardware occurred in two parts of the country: Florida, where the This performance analysis provided Advancements in technology that hardware was retrieved; and Utah, a wealth of valuable information occurred during the decades-long where it underwent in-depth inspection and created a synergy to drive program were a further source of and refurbishment. On recovery, a team improvements in motor performance, change. Engineers incorporated new of 15 motor engineers conducted what implemented through motor technologies into motor design and was termed an “open assessment,” manufacturing and processing. processing as the technology could be primarily focusing on exterior This recovery and postflight capability proven. Incorporating braided carbon components. After retrieval, teams of was particularly important for the fiber material as a thermal barrier in the specialists rigorously dissected, long-term Space Shuttle Program since, nozzle-to-case joint is one example. measured, sampled, and assessed joints,
168 Engineering Innovations Field Joint Comp arison for Use on Reusable Solid Rocket Motor
Cork Insulation
Temperature Sensor (added)
Thermal Barrier Leak Check Port Heater and Heat Relocated and Modi ed Transfer Cement (added) Leak V-2 Filler (added) Check Fluorocarbon Kevlar® Retainer Strap Port Primary O-ring Interference Fit (added) Vent Port in Front of Zinc Grease Primary O-ring (added) Capture Feature Chromate Bead O-ring (added) Putty Custom Shims (added) Capture Feature (added) Shim Resin Technology 455 J-joint Pressurization Cork Slot (added) Fluorocarbon Insulation Secondary Larger Pressure-sensitive O-ring Longer Pins Grooves Adhesive (added) and New and O-ring Size Nonvented Joint Filled Retention Band (added) Insulation (added) Cork Insulation Insulation Gap Resin Technology 455
High-performance M otor Reus able S olid Rock et Motor . d e v
Reusable solid rocket motors incorporated significant improvements over the earlier shuttle motors in the design of the joints between the r e s e
main segments. Redesign of this key feature was part of the intensive engineering redesign and demonstration feat accomplished following r
s t
the Challenger accident. The result was a fail-safe joint/seal configuration that, with continued refinement, had a high demonstrated reliability. h g i r
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Each joint, with its redundant seals and multiple redundant seal protection features, could be pressure checked during assembly to ensure A
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a good pressure seal was achieved. A similar design approach was implemented on the igniter joints during that same time period. T A
© bondlines, ablatives, fasteners, and data were fed into a system-wide The postflight analysis program virtually all remaining flight hardware. database containing documentation collected the actual flight performance Engineers promptly evaluated any dating back to the program’s inception. data—most of which would not have significant observations that could The wealth of information available been available if the motors had not affect the orbiting vehicle or the next for performance trend analysis was been recovered. motor launch sets. unmatched by any other solid rocket Through this tightly defined process, motor manufacturing process in the Before the motor was returned to the engineers were able to address the world. Gates and checks within the flight inventory, the recovered metal subtle effects that are often a result of an system ensured the full investigation parts were inspected for corrosion, unintended drift in the manufacturing of any anomalies to pinpoint root deformations, cracks, and other potential process or new manufacturing materials cause and initiate corrective action. damage. Dimensional measurement introduced into the process. The
Engineering Innovations 169 process addressed these concerns in the incipient phase rather than allowing for a potentially serious issue to escalate undetected. The ultimate intangible benefit of this program was greater reliability, as demonstrated by the following two examples. Postflight assessment of nozzle bondlines was a catalyst to augment adhesive bonding technology and . d
substantially improve hardware quality e v r e s
and reliability. Storage controls for e r
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epoxy adhesives were established g i r
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in-house and with adhesive suppliers. . K T A Surface preparation, cleanliness, © adhesive primer, and process timelines In Utah, rigorous test program included 53 reusable solid rocket motor ground tests between 1977 and were established. Adhesive bond 2010. Spectators flocked by the thousands to witness firsthand the equivalent of 15 million horsepower quality and robustness were increased safely unleashed from a vantage point of 2 to 3 km (1 to 2 miles) away. by an order of magnitude.
Postflight inspections also occasionally the highest levels of dependability and included a suite of sensors not limited revealed gas paths through the safety for the hardware. Immediate to accelerometers, pressure transducers, nozzle-to-case joint polysulfide thermal challenges posed by the 570-metric-ton calorimeters, strain gauges, barrier that led to hot gas impingement (1 .2-million-pound) motor included thermocouples, and microphones. on the wiper O-ring—a structure handling, tooling, and developing a Beyond overall system assessment and protecting the primary O-ring from 17.8-meganewton (4,000,000-pound- component qualification, benefits of thermal damage. While this condition force) thrust-capable ground test stand; full-scale testing included the did not pose a flight risk, it did indicate and designing a 1,000-channel data opportunity to enhance engineering performance failed to meet design handling system as well as new support expertise and predictive skills, improve intent. The root cause: a design that systems, instrumentation capability, data engineering techniques, and conduct was impossible to manufacture acquisition, and countdown procedures. precise margin testing. The ability to perfectly every time. Engineers tightly measure margins for many Hot-fire testing of full-scale rocket resolved this concern by implementing motor process, material, components, motors in the Utah desert became a a nozzle-to-case joint J-leg design and design parameters provided hallmark of the reusable solid rocket similar to that successfully used on valuable verification data to motor development and sustainment case field joints and igniters. demonstrate whether even the slightest program. Individual motor rockets modification was safe for flight. were fired horizontally, typically Robust Systems Testing once or twice a year, lighting up the Quick-look data revealed basic ballistics mountainside with the brightness of a performance—pressure and thrust The adage “test before you fly,” blazing sun, even in broad daylight. measurements—that could be compared adopted by the Space Shuttle Program, with predicted performance and historic was the standard for many reusable Following a test firing, quick-look data data for an initial assessment. solid rocket motor processes and were available within hours. Full data material, hardware, and design changes. analyses required several months. Full analysis included scrutiny of all What ATK, the manufacturer, was able data recorded during the actual test as On average, NASA collected between to learn from the vast range of data well as additional data gathered from 400 and 700 channels of data for each collected and processed through visual inspections and measurements test. Instrumentation varied according preflight and ground testing ensured of disassembled hardware, similar to test requirements but typically
170 Engineering Innovations to that of postflight inspection. Noteworthy elements of the motor Orbital Propulsion Engineers assessed specific data tied process control program included to test objectives. When qualifying a an extensive chemical fingerprinting Systems— new motor insulation, for example, program to analyze and monitor the Unique Development posttest inspection would additionally quality of vendor-supplied materials, include measurements of remaining the use of statistical process control to Challenges insulation material to calculate the rate better monitor conditions, and the of material loss. comprehensive use of witness Until the development of the Space panels—product samples captured from Shuttle, all space vehicle propulsion Subscale propellant batch ballistics the live manufacturing process and systems were expendable. Influenced tests, environmental conditioning analyzed to validate product quality. by advances in technologies and testing, vibration tests, and custom materials, NASA decided to develop sensor development and data With scrupulous process control, a reusable propulsion system. acquisition were also successful ATK and NASA achieved an even Although reusability saved overall components of the program to provide greater level of understanding of the costs, maintenance and turnaround specific reliability data. materials and processes involved with costs offset some of those benefits. reusable solid rocket motor processing. As a result, product output became NASA established a general redundancy Culture of Continual more consistent over the life of the requirement of fail operational/fail safe Improvement program. Additionally, partnerships for these critical systems: Orbital Maneuvering System, Reaction Control The drive to achieve 100% mission with vendors and suppliers were System, and Auxiliary Power Unit. success, paired with the innovations of strengthened as increased performance In addition, engineers designed the pre- and postflight testing that allowed measurement and data sharing created propulsion systems for a life of 100 performance to be precisely quantified, a win-win situation. missions or 10 years combined storage resulted in an operating culture in and operations. Limited refurbishment which the bar was continually raised. An Enduring Legacy was permitted at the expense of higher Design and processing improvements The reusable solid rocket motor was operational costs. were identified, pursued, and more than an exceptional rocket that implemented through the end of safely carried astronauts and hundreds Orbital Maneuvering System the program to incrementally reduce of metric tons of hardware into orbit risk and waste. Examples of relatively for more than 25 years. Throughout the The Orbital Maneuvering System late program innovations included: Reusable Solid Rocket Motor Program, provided propulsion for the Orbiter permeable carbon fiber rope as a engineers and scientists generated the during orbit insertion, orbit thermal protection element in various technical know-how in design, test, circularization, orbit transfer, nozzle and nozzle/case joints; analysis, production, and process rendezvous, and deorbit. NASA faced structurally optimized bolted joints; control that is essential to continued a major challenge in selecting the reduced stress forward-grain fin space exploration. The legacy of the propellant. The agency originally chose transition configuration; and improved first human-rated reusable solid rocket liquid oxygen and liquid hydrogen adhesive bonding systems. motor will carry on in future decades. propellants. However, internal volume This culture, firmly rooted in the wake In the pages of history, the shuttle constraints could not be met for a of the Challenger accident, led to a reusable solid rocket motor will be vehicle configuration that provided a comprehensive process control program known as more than a stepping-stone. payload of 22,680 kg (50,000 pounds) with systems and tools to ensure It will also be regarded as a benchmark in a bay measuring 4.6 m (1 5 ft) in processes were appropriately defined, by which future solid-propulsion diameter and 18.3 m (60 ft) in length. correctly performed, and adequately systems will be measured. This, coupled with concerns regarding maintained to guarantee reliable and complexity of cryogenic propellants, repeatable product performance. led to the consideration of storable hypergolic propellants.
Engineering Innovations 171 Thus, NASA adopted an interconnect system in which the Reaction Control Orbital Mane uvering System/React ion Control System System used Orbital Maneuvering System propellants because of cost, weight, and lower development risk. Helium Orbital Maneuvering TanksTanks System Engine Disadvantages of a storable propellant system were higher maintenance Fuel TTankank Primary requirements resulting from their Thrusters (12 total) corrosive nature and hazards to Fuel TankTaank personnel exposed to the toxic propellants. NASA partially addressed VVerVernierernier Thrusters (2 total) these considerations by incorporating the Orbital Maneuvering System into a removable modular pod. This allowed maintenance and refurbishment Helium TankTank of those components exposed to hypergols to be separated from other Oxidizer TankTanank turnaround activities. Oxidizerdizer TankTank = Reaction ControlControl System = Orbital Maneuvering System For ground operations, it was not practical to remove modules for each turnaround activity, and sophisticated equipment and processes were required for servicing between flights. Fluid and gas connections to the propellants and pressurants used quick Orbital Maneuvering disconnects to allow servicing on System/Reaction the launch pad, in Orbiter processing Control System facilities, and in the hypergolic pods viewed from maintenance facility. However, quick the underside. disconnects occasionally caused problems, including leakage that damaged Orbiter thermal tiles. Engineers tested and evaluated many ground support equipment design NASA ultimately selected Modular Design Presents concepts at the White Sands Test monomethylhydrazine as the fuel Obstacles for Ground Support Facility (WSTF). In particular, they and nitrogen tetroxide as the oxidizer tested, designed, and built the Trade studies and design approach for this system. As these propellants equipment used to test and evaluate investigations identified challenges were hypergolic—they ignited when the propellant acquisition screens inside and solutions. For instance, cost and coming into contact with each the propellant tanks before shipment to weight could be reduced with a other—no ignition device was Kennedy Space Center for use on flight common integrated structure for the needed. Both propellants remained vehicles. The Orbital Maneuvering Orbital Maneuvering System and liquid at the temperatures normally System/Reaction Control System Fleet Reaction Control System. This experienced during a mission. Leader Program used existing integrated structure was combined with Electrical heaters prevented freezing qualification test articles to detect and the selection of nitrogen tetroxide and during long periods in orbit when the evaluate “life-dependent” problems monomethylhydrazine propellants. system was not in use. before these problems affected the
172 Engineering Innovations shuttle fleet. This program provided a test bed for developing and evaluating ground support equipment design changes and improving processes and procedures. An example of this was Henry Pohl the Reaction Control System Thruster Director of Engineering at Johnson Space Center Purge System, which used low-pressure (1986-1993). nitrogen to prevent propellant vapors from accumulating in the thruster “To begin to understand the challenges of chamber. This WSTF-developed operating without gravity, imagine removing ground support system proved the commode from your bathroom floor, bolting it to the ceiling. And then try beneficial in reducing the number of in-flight thruster failures. to use it. You would then have a measure of the challenges facing NASA.”
Additional Challenges Stable combustion was a concern for cool and therefore less subject to failure eliminated mechanical manufacturing NASA. In fact, stable combustion has from either burnout or thermal cycling. errors and increased injector life and always been the most expensive combustion efficiency. schedule-constraining development To accomplish precise injector issue in rocket development. For the fabrication, engineers implemented The combustion chamber was Orbital Maneuvering System engine, platelet configuration. The fuel and regenerative-cooled by fuel flowing engineers investigated injector pattern oxidizer flowed through the injector in a single pass through non-tubular designs combined with acoustic cavity and impinged on each other, causing coolant channels. The chamber was concepts. In propulsion applications mixing and combustion. Platelet composed of a stainless-steel liner, an with requirements for long-duration technology, consisting of a series of electroformed nickel shell, and an aft firings and reusability, cavities had an thin plates manufactured by photo flange and fuel inlet manifold assembly. advantage because they were easy to etching and diffusion bonded together, Its structural design was based on life
Formation of Metal Nitrates Caused Valve Leaks
Being the first reusable spacecraft—and in particular, the first to Subsequent valve cycling caused damage to the Teflon ® valve seat, use hypergolic propellants—the shuttle presented technical further exacerbating the leakage until sufficient nitrate deposition challenges, including leaky and sticky propellant valves in the resulted in “gumming” up the valve. At that point, the valve was Reaction Control System thrusters. Early in the program, failures in either slow to operate or failed to operate. this system were either an oxidizer valve leak or failure to reach full Multiple changes reduced the metal nitrate problem but may have chamber pressure within an acceptable amount of time after the contributed to fuel valve seat extrusion, which manifested years thruster was commanded on. NASA attributed both problems to the later. The fuel valve extrusion was largely attributed to the use buildup of metal nitrates on and around the valve-sealing surfaces. of throat plugs. These plugs trapped oxidizer vapor leakage in the Metal nitrates were products of iron dissolved in the oxidizer combustion chamber, which subsequently reacted at a low level when purchased and iron and nickel that were leached out of the of fuel that had permeated the Teflon ® fuel valve seat. This problem ground and flight fluid systems. When the oxidizer was exposed was successfully addressed with the implementation of the to reduced pressure or allowed to evaporate, metal nitrates NASA-developed thruster nitrogen purge system, which kept the precipitated out of solution and contaminated the valve seat. thruster combustion chamber relatively free of propellant vapors.
Engineering Innovations 173 cycle requirements, mechanical loads, thrust and aerodynamic loading on An Ordinary Solution to the Extraordinary the nozzle, ease of fabrication, and Challenge of Rain Protection weight requirements. The nozzle extension was radiation During operations, Orbiter cooled and constructed of columbium engines needed rain metal consistent with experience gained during the Apollo Program. The protection after the mounting flange consisted of a bolt ring, protective structure was made from a forging and a tapered moved away and protective section, that could either be spun or ground covers were made from a forging. The forward and removed. This requirement aft sections were made from two panels each. This assembly was bulge formed protected the three to the final configuration and the upward-facing engines stiffening rings were attached by and eight of the left-side welding. The oxidation barrier diffusion engines from rainwater operation was done after machining accumulation on the launch was completed. pad. The up-firing engine A basic design challenge for the covers had to prevent bipropellant valve was the modular valve. The primary aspect of the water accumulation that Tyvek ® covers shown installed assembly design was modularization, could freeze in the injector on forward Reaction Control System thrusters (top) and a which reduced fabrication problems passages during ascent. typical cover (right). Note that and development time and allowed The side-firing engine the covers were designed to servicing and maintenance goals to be fit certain thruster exit plane covers prevented water met with lower inventory. configurations. from accumulating in NASA Seeks Options as the bottom of the chamber and protected the chamber pressure sensing ports. Costs Increase Freezing of accumulated water during ascent could block the sensing port and The most significant lesson learned cause the engine to be declared “failed off” when first used. The original design during Orbital Maneuvering System ® concept allowed for Teflon plugs installed in the engine throats and a combination development was the advantage of of Teflon ® plugs tied to a Teflon ® plate that covered the nozzle exit. This concept developing critical technologies before added vehicle weight, required special procedures to eject the plugs in flight, and initiating full-scale hardware designs. risked accidental ejection in ascent that could damage tiles. The solution used The successful completion of predevelopment studies not only ordinary plastic-coated freezer paper cut to fit the exit plane of the nozzle. Tests reduced total costs, also it minimized proved this concept could provide a reliable seal under all expected rain and wind schedule delays. conditions. The covers were low cost, simple, and added no significant weight. In the 1980s, NASA began looking The thruster rain cover material was changed to Tyvek ® when NASA discovered for ways to decrease the cost of pieces of liberated plastic-coated paper beneath the cockpit window pressure component refurbishment and repair. seals. The new Tyvek ® covers were designed to release at relatively low vehicle NASA consolidated engineering, velocity so that the liberated covers did not cause impact damage to windows, evaluation, and repair capabilities for tile, or any other Orbiter surface. many components, and reduced overall costs. Technicians serviced, acceptance
174 Engineering Innovations
(2 total) Checkout Panel
Thruster
Purge and (2 total)
Vernier
ank T
anel P
elium H
Fuel Tank ervicing S
Tank
Oxidizer
Disconnect Panel
lectrical E
Primary Thrusters (14 total)
Forward Reaction Control System
Forward Reaction Control System on Discovery.
tested, and prepared all hypergolic requirement of a fail-operational/fail- propellant tank acquisition system wetted components for reinstallation safe design introduced complexity of design because of changes in the on the vehicles. additional hardware and a complex gravitational environment. critical redundancy management system. The reuse requirement posed problems NASA Makes Effective Selections Reaction Control System in material selection and compatibility, As with the Orbital Maneuvering ground handling and turnaround The Reaction Control System provided System, propellant selection was procedures, and classical wear-out propulsive forces to control the motion important for the Reaction Control problems. The requirement for both of the Orbiter for attitude control, System. NASA chose a bipropellant of on-orbit operations and re-entry into rotational maneuvers, and small velocity monomethylhydrazine and nitrogen changes along the Orbiter axes. The Earth’s atmosphere complicated
Low Temperatures, Increased Leakage, and a Calculated Solution
Some primary thruster valves could indicated that tetrafluoroethylene Teflon ® To reduce susceptibility to cold leakage, leak when subjected to low temperature. underwent a marked change in the engineers machined Teflon ® at 0°C (32°F) NASA discovered this problem when thermal expansion rate in a designated to ensure uniform dimensions with they observed liquid dripping from the temperature range. Because machining, adequate seat material exposed at system level engines during a cold done as a part of seat fabrication, was reduced temperatures and raised the environment test. The leakage became accomplished in this temperature range, thruster heater set points to maintain valve progressively worse with increased some parts had insufficient seat material temperature above 16°C (60°F). cycling. Continued investigation exposed at reduced temperatures.
Engineering Innovations 175 Cracks Prompt Ultrasonic Inspection
Late in the Space Shuttle Program, NASA discovered cracks in a thruster injector. The thruster was being refurbished at Reaction Control System Valve Module White Sands Test Facility (WSTF) during Primary Thruster the post-Columbia accident Return to Flight time period. The cracks were Injector markedly similar to those that had occurred Combustion in injectors in 1979 and again in 1982. Chamber
These earlier cracks were discovered during manufacturing of the thrusters and occurred during the nozzle insulation bake-out process. Results from the Columbium Nozzle laboratory testing indicated that cracks were developed due to chemical processing and manufacturing. In addition to using leak testing to Oxidizer Tube screen for injector cracking, NASA Fuel Tube Acoustic Cavity engineers developed and implemented Mounting Flange an ultrasonic inspection procedure to screen for cracks that measured less Injector Flange Relief than the injector wall thickness. Injector Radius Mounting Bolt Cracks Injector Face The marked similarity of the crack location Chamber Wall and crack surface appearance strongly Cracked Relief Radius suggested the WSTF-discovered cracks
were due to the original equipment Relief Radius manufacturing process and were not flight induced or propagated. Laboratory Cracks tests and analyses confirmed that those cracks were induced in manufacturing. The cracks had not grown significantly over the years of the thruster’s use and its Reaction Control System thruster cross sections showing the crack location and its actual many engine firings. Laboratory surface appearance. nondestructive testing showed that the escape detection and cracked thrusters cracks due to the service environment was original ultrasonic inspection process was could have been placed in service. The fact a significant factor in the development of not very reliable and it was possible that that there was no evidence of crack growth flight rationale for the thrusters. manufacturing-induced cracks could associated with the WSTF-discovered
176 Engineering Innovations tetroxide system, which allowed for a program to determine the parameters steam generated was vented overboard. integration of this system with the that caused the iron nitrate formation Use of this system enabled restarts at Orbital Maneuvering System. This and implement procedures to prevent any time after the cooling process, propellant combination offered a its formation in the future. This which required a 21 0-second delay. favorable weight tradeoff, reasonable resulted in understanding the development cost, and minimal relationship between iron, water, Improved Machining and development risk. nitric oxide content, and nitrate Manufacturing Solves Valve Issue formation. The agency developed NASA selected a screen tank as a Development of a reliable valve production and storage controls as reusable propellant supply system to to control fuel flow into the gas well as filtration techniques to provide gas-free propellants to the generator proved to be one of the remove the iron, which resolved the thrusters. Screen tanks worked by using most daunting tasks of the propulsion iron nitrate problem. the surface tension of the liquid to form systems. The valve was required to a barrier to the pressurant gas. The pulse fuel into the gas generator at propellant acquisition device was made Auxiliary Power Unit frequencies of 1 to 3 hertz. Problems of channels covered with a finely woven with the valve centered on leakage and steel mesh screen. Contact with liquid The Auxiliary Power Unit generated limited life due to wear and breakage wetted the screen and surface tension of power to drive hydraulic pumps that of the tungsten carbide seat. NASA’s the liquid prevented the passage of gas. produced pressure for actuators to considerable effort in redesigning the The strength of the liquid barrier was control the main engines, aero surfaces, seat and developing manufacturing finite. The pressure differential at which landing gear, brakes, and nose wheel processes resulted in an intricate seat gas would be forced through the wetted steering. The Auxiliary Power Unit design with concentric dual sealing screen was called the “bubble point.” shared common hardware and systems surfaces and redesigned internal flow When the bubble point was exceeded, with the Hydraulic Power Unit used passages. The seat was diamond-slurry the screen broke down and gas was on the solid rocket motors. The shuttle honed as part of the manufacturing transferred. If the pressure differential needed a hydraulic power unit that process to remove the recast layer left was less than the bubble point, gas could operate from zero to three times by the electro-discharge machining. could not penetrate the liquid barrier gravity, at vacuum and sea-level This recast layer was a source of and only liquid was pulled through pressures, from -54°C to 107°C stress risers and was considered the channels. NASA achieved their (-65°F to 225°F), and be capable of one of the primary factors causing goal in designing the tank to minimize restarting. NASA took the basic seat failure. The improved design the pressure loss while maximizing the approach of using a small, high-speed, and machining and manufacturing amount of propellant expelled. monopropellant-fuel, turbine-powered processes were successful. unit to drive a conventional aircraft- Several Reaction Control System type hydraulic pump. Additional Challenges and component failures were related to Subsequent Solutions nitrate contamination. Storage of If the Auxiliary Power Unit was oxidizer in tanks and plumbing that restarted before the injector cooled to During development testing of the contained iron caused contamination less than 204°C to 232°C (400°F to gear box, engineers determined in the propellant. This contamination 450°F), the fuel would thermally that the oil pump may not funtion formed a nitrate that could cause valve decompose behind the injector panels satisfactorily on orbit due to low leakage, filter blockage, and and damage the injector and the Gas pressure. It became necessary to interference in sliding fits. The most Generator Valve Module. Limited provide a fluid for the pump to displace prominent incident was the failure of hot-restart capability was achieved by to assure the presence of oil at the a ground half-quick disconnect to adding an active water cooling system inlet and to have a mechanism to close, resulting in an oxidizer spill on to the gas generator to be used only for provide needed minimum pressure at the launch pad. NASA implemented hot restarts. This system injected water startup and during operation. into a cavity within the injector. The
Engineering Innovations 177 The Auxiliary Power Unit was provided safety features that would inspections. With these controls, designed with a turbine wheel radial allow operation within the existing results of fracture mechanics analyses containment ring and a blade tip seal degree of containment. The agency used showed the theoretical life to be 10 and rub ring to safely control failures an over-speed safety circuit to times the 100-mission requirement. of the high-speed assembly. The automatically shut down a unit at 93,000 With these improvements, the Auxiliary containment ring was intended to keep revolutions per minute. To provide Power Unit demonstrated success of any wheel fragments from leaving the further insurance against wheel failure, design and exhibited proven durability, Auxiliary Power Unit envelope. NASA NASA imposed stringent flaw detection performance, and reusability.
NASA Encounters Obstacle Course in Turbine Wheel Design
The space agency faced multiple challenges with the development The shroud cracking problem was related to material selection of the turbine wheel. Aerodynamically induced high-cycle fatigue and the welding process. Increased strength and weld caused cracking. Analysis indicated this part of the blade could be characteristics were achieved by changing the shroud material. removed with a small chamfer at the blade tip without significant Engineers developed a controlled electron beam weld procedure effect on performance. This cracking problem was resolved by to ensure no overheating of the shroud. These actions eliminated careful design and control of electromechanical machining. the cracking problem.
Bearing Housing Drive Spur Lube Sleeve Gear Straight Pin Bolt
Ball Bearing
Turbine Blade
Blade Root Cracking
178 Engineering Innovations Stress Corrosion and Propellant Ignition
One of the most significant Auxiliary Power overheated the units, causing the residual hydrazine in the catalyst bed after Auxiliary Unit problems occurred during the STS-9 hydrazine to detonate after landing. The Power Unit shutdown. (1983) mission when two of the three units fire investigation determined the source of Initial corrective actions included removal caught fire and detonated. Postflight the leaks to be nearly identical cracks in of the electrical machined recast layer analysis indicated the presence of the gas generator injector tubes in both on the tube inside diameter and an hydrazine leaks in Auxiliary Power Units 1 units. Laboratory tests further determined improved assembly of the injector tube. and 2 when they were started for re-entry that the injector tube cracks were due to Later, resistance to stress corrosion and while still in orbit. The leaking hydrazine stress corrosion from ammonium hydroxide general corrosion was further improved by subsequently ignited and the resulting fire vapors generated by decomposition of chromizing the injector tubes.
Catalyst
O-ring Grooves Injector Stem Hot Gas Liquid Out To Hydrazine Turbine In Wheel
STS-9 Cracked Injector Stem
Summary technologies required to meet encountered. In addition, several changing requirements, and continued problems were not anticipated. NASA The evolution of orbital propulsion with improvements based on flight met these challenges, as demonstrated systems for the Space Shuttle experience. The design requirements by the success of these systems. Program began with Apollo Program for 100 missions, 10 years, and reuse concepts, expanded with new presented challenges not previously
Engineering Innovations 179 Propulsion Systems Pioneering and Hazardous Inspection Tool Gas Detection Contamination Scanning Shuttle propulsion had hazardous gases of Bond Surfaces requiring development of detection systems including purged compartments. Bonding thermal insulation to metal This development was based on lessons case surfaces was a critical process learned from the system first used during in solid rocket motor manufacturing Saturn I launches. during the Space Shuttle Program. NASA performed an exhaustive review Surfaces had to be immaculately Inspection technology capitalizing on the of all available online monitoring photoelectric effect provided significant benefits clean for proper adherence. The steel mass spectrometry technology for the over the traditional method of visual inspection shuttle. The system the agency selected alloy was susceptible to corrosion using handheld black lights. The technology was for the prototype Hazardous Gas and was coated with grease for developed through a NASA/industry partnership managed by Marshall Space Flight Center. Detection System had an automated protection during storage. That Specific benefits included increased accuracy in high-vacuum system, a built-in grease, and the solvents to remove it, contamination detection and an electronic data computer control interface, and the record for each hardware inspection. became potential contaminants. ability to meet all program-anticipated detection limit requirements. The improvement of contamination inspection techniques was initiated in the late The instrument arrived at Kennedy 1980s. The development of a quantitative and recordable inspection technique was Space Center (KSC) in December 1975 based on the physics of optically stimulated electron emission (photoelectric effect) and was integrated into the sample technology being developed at NASA’s Marshall Space Flight Center at the time. delivery subsystem, the control and data subsystem, and the remote Fundamentally, incident ultraviolet light excites and frees electrons from the metal control subsystem designed by KSC. surface. The freed electrons having a negative charge are attracted to a positively Engineers extensively tested the charged collector ring in the “Con Scan” (short for Contamination Scanning) sensor. unit for functionality, detection limits When contamination exists on a metal surface, the amount of ultraviolet radiation and dynamic range, long-term drift, that reaches the surface is reduced. In turn, the current is reduced, confirming the and other typical instrumental presence of a contaminant. performance characteristics. In May 1977, KSC shipped the prototype Approximately 90% of each reusable solid rocket motor barrel assembly was Hazardous Gas Detection System to inspected using automated Con Scan before bond operations. Technicians mounted Stennis Space Center to support the shuttle main propulsion test article the sensor on a robotic arm, which allowed longitudinal translation of the sensor as engine test firings. The system the barrel assembly rotated on a turntable. Inspection results were mapped, showing remained in use at Stennis Space color-coded contamination levels (measured current) vs. axial and circumferential Center for 12 years and supported the locations on the case inner diameter. Color coding made acceptable and rejected testing of upgraded engines. areas visually apparent. The first operational Hazardous Gas By pioneering optically stimulated electron emission technology, which was Detection System was installed for the system on the Mobile Launch engineered into a baseline inspection tool, the Space Shuttle Program significantly improved contamination control methods for critical bonding applications.
180 Engineering Innovations Platform-1 during the late summer Originally, NASA declined to provide During launch countdown, NASA of 1979. Checkout and operations redundancy for the Hazardous Gas detected the aft fuselage hydrogen procedure development and activation Detection System due to a lack of a leak. It was then apparent that STS-35 required almost 1 year, but the system launch-on-time requirement; however, had experienced two separate leaks. was ready to support initial purge the agency subsequently decided that The Space Shuttle Program director activation and propellant loading tests redundancy was required. After a appointed a special tiger team to in late 1980. A special test in which detailed engineering analysis followed investigate the leak problem. This team engineers introduced simulated leaks by lab testing of candidate mass suspected that the Hazardous Gas of hydrogen and oxygen into the spectrometers, the space agency Detection System was giving Orbiter payload bay, lower midbody, selected the PerkinElmer MGA -1 200 erroneous data, and brought aft fuselage, and the External Tank as the basis of the backup Hazardous 10 experts from Marshall Space Flight intertank area represented a significant Gas Detection System. This backup Center to assess the system design. milestone. The system accurately was an ion-pumped, magnetic-sector, KSC design engineering provided an detected and measured gas leaks. multiple-collector mass spectrometer in-depth, 2-week description of the widely used in operating rooms and design and performance details of both After the new system’s activation issues industrial plants. Although the first the Hazardous Gas Detection System were worked out, it could detect and systems were delivered in late 1985, and the backup system. The most measure small leaks from the Main full installation on all mobile launch compelling evidence of the validity of Propulsion System. The Hazardous platforms did not occur until NASA the readings was that both systems, Gas Detection System did not become completed the Return to Flight which used different technology, had visible until Space Transportation activities following the Challenger measured identical data, and both System (STS)-6—the first launch of accident in 1986. systems had recorded accurate the new Orbiter Challenger—during a calibration data before and after flight readiness test. In this test, the In May 1990, the Hazardous Gas leakage detection. After a series of countdown would proceed normally Detection System gained attention mini-tanking tests—each with to launch time, the Orbiter main engines once again when NASA detected a increased temporary instrumentation— would ignite, but the Solid Rocket hydrogen leak in the Orbiter aft engineers located and repaired the leak, Booster engines would not ignite and fuselage on STS-35. The space agency and STS-35 lifted off for a successful the shuttle would remain bolted to the also detected a hydrogen leak at the mission on December 9, 1990. launch pad during a 20-second firing of External Tank to Orbiter hydrogen the main engines. The STS-1 firing test umbilical disconnect and thought that The Hazardous Gas Detection System for Columbia had proceeded normally, the aft fuselage leakage indication was and backup Hazardous Gas Detection but during Challenger’s firing test, the due to hydrogen from the external leak System continued to serve the Hazardous Gas Detection System migrating inside the Orbiter. Workers shuttle until 20 01, when both systems detected a leak exceeding 4,000 parts rolled STS-35 back into the Vertical were replaced with Hazardous Gas per million. Rerunning the firing test Assembly Building and replaced the Detection System 2000—a modern and performing further leak hunting umbilical disconnect. Meanwhile, state-of-the-art system with a common and analysis revealed a number of STS-38 had been rolled to the pad and sampling system and identical twin faults in the main engines. The manager leakage was again detected at the quadrupole mass spectrometers for shuttle operation propulsion umbilical disconnect, but not in the aft from Stanford Research Institute. stated that all the money spent on the fuselage. STS-38 was also rolled back, The Hazardous Gas Detection System Hazardous Gas Detection System, and its umbilical disconnect was served for 22 years and the backup and all that would ever be spent, was replaced. The ensuing investigation Hazardous Gas Detection System paid for in those 20 seconds when the revealed that manufacturing defects in served for 15 years. leak was detected. both units caused the leaks, but not before STS-35 was back on the pad.
Engineering Innovations 181