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Review of flaws and damages in space launch vehicle: Structures

Article  in  Journal of Intelligent Material Systems and Structures · January 2013

DOI: 10.1177/1045389X12458041

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Review article

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Review of flaws and damages in spacelaunch vehicle: Structures

Jung-Ryul Lee and Dipesh Dhital

AbstractA space launch vehicle is a complex engineering structure. It consists of several structural and system components thatare stably stored for extended periods and that at the launch moment, must function precisely. However, this presents anumber of unique challenges as well as the possibility of various flaws and damages during the manufacturing, assembly,or ground handling phase. Mechanical and chemical complexity, long service lives, aging materials, and designs with smallmargins are typical of space launch vehicle components. The performance and characteristics of the components can besignificantly affected by degradation resulting from such flaws and damages, and this degradation in turn might lead to fail-ure of the entire space mission. Damage detection at the earliest possible stage enables implementation of possible pre-ventive measures, and evaluation of component reliability and readiness either at the manufacturing level or during fieldinspections. This review study lists such possible flaws/damages for various space launch vehicle components. This infor-mation could be beneficial for determining and developing reliable nondestructive evaluation and structural health moni-toring techniques.

KeywordsSpace launch vehicle, flaws and damages, structural health monitoring, nondestructive evaluation

Introduction

A space launch vehicle or a carrier rocket carries pay-load or human crew into outer space. They have beenused to send manned spacecrafts, unmanned spaceprobes, and satellites into space. Soyuz and Protonlaunchers of Russia, the Ariane series of Europe, theSpace Shuttle, Atlas, Delta, and Titan families of vehi-cles of the United States are few examples (Gregersen,2009). They are generally categorized as either expend-able launch vehicle (ELV) or reusable launch vehicle(RLV). Most space probes get launched on expendable‘‘throw away’’ rockets and are designed for a single use.The only partially reusable systems that have ever beenemployed are the US Space Shuttle and the SovietBuran vehicle (which was flown only once) (Gregersen,2009; Pelt, 2006). The Space Shuttle consists of an orbi-ter, a large external tank (ET), and two enormous solidrocket boosters. It returns astronauts and materials tothe earth and is partially reusable, although an enor-mous effort is needed to refurbish the vehicle after eachflight. The reason that expendable rockets are still usedmore often than the reusable systems such as the SpaceShuttle is that they are much less expensive to developand much easier to operate. In fact, the maintenance ofthe Space Shuttle is so costly that for the relatively few

times per year that a satellite needs to be launched, it ischeaper to build and use an expendable, one-shotrocket (Pelt, 2006). Attempts to develop a fully RLV toreduce the cost of access to space have so far not beensuccessful, primarily because the propulsion system andmaterials needed for successful development of such avehicle have not been available (Gregersen, 2009).

An ELV in general comprises one or more rocketengines; fuel tanks carrying the fuel for those engines;guidance, navigation, and control systems; a payload;and a structure housing all of these elements, to whichextra engines may be attached for added lift(Gregersen, 2009), as shown in Figure 1 (Pelt, 2006).The basic approach to launch vehicle design is to dividethe vehicle into stages. Most ELVs in use today haveonly two or three stages to reach its orbital velocity,but in the past, had up to five stages. The first stage,

Department of Aerospace Engineering and LANL-CBNU Engineering

Institute Korea, Chonbuk National University, Deokjin-gu, Jeonju,

Republic of Korea

Corresponding author:

Jung-Ryul Lee, Department of Aerospace Engineering and LANL-CBNU

Engineering Institute Korea, Chonbuk National University, 567 Baekje-

daero, Deokjin-gu, Jeonju 561-756, Republic of Korea.

Email: leejrr@jbnu.ac.kr

which is the heaviest part of the vehicle, has the largestrocket engines and the largest fuel and oxidizer tanksand provides the highest thrust to the vehicle; its task isto impart the initial thrust needed to overcome earth’sgravity and thus to lift the total weight of the vehicleand its payload off the earth. When the first-stage pro-pellants are used up, this stage is detached from theremaining parts of the launch vehicle and falls back tothe earth. With the weight of the first stage gone, thesecond stage, with its own rocket engines and propel-lants, continues to accelerate the vehicle. A particularlaunch vehicle can be configured in several differentways, depending on its mission and the weight of thespacecraft to be launched, by adding various numbersof strap-on boosters, usually solid rocket motors(SRMs), to the vehicle’s first stage or using differentupper stages (Gregersen, 2009).

In theory, launch vehicles can be thermal-propelled(chemical, nuclear, solar, and laser), electric-propelled,or nuclear-propelled (radioisotype and explosion) sys-tems. However, the primary propulsion system forlaunch vehicles is restricted currently to solid-propelledor liquid-propelled chemical rockets (Fortescue et al.,

2003). Solid systems are usually called motors, andliquid systems are referred to as engines (Gregersen,2009). The burning of chemical propellants, solid orliquid, at high pressure liberates large quantities ofenergy from a compact volume. The subsequentexpansion of these high-temperature products ofcombustion through a nozzle converts thermal energyinto directed kinetic energy for rocket propulsion(Fortescue et al., 2003). Most of the weight of thelaunch vehicle is actually the combined weights of itspropellants, the fuel, and the oxidizer needed to burnthe fuel, which often makes up 80% or more of thetotal weight of a launch vehicle–payload combinationprior to launch (Gregersen, 2009).

This review study is about the flaws and damages inspace launch vehicle structures. It lists the possible pat-terns of damages that can be considered as the hot spotlocations in different structures like payload fairing(sandwich structure and grid-stiffened structures),rocket propellant tanks/fuel tanks, payload adapter,and thermal protection system (TPS). All of these com-ponents are reviewed in detail in the following chapters.

Space launch—failures and causes

A space launch failure is an unsuccessful attempt toplace a payload into its intended orbit. This definitionincludes all catastrophic launch mishaps involvinglaunch vehicle destruction or explosion, significantreduction in payload service life, and extensive effort orsubstantial cost for mission recovery (Chang, 2005).Launch vehicles are designed to be as light as possibleto maximize their payload lifting capability. As a result,every part in a launch vehicle operates close to itsbreaking point during a launch, as the vehicle under-goes the stresses associated with accelerating past thespeed of sound and transiting the atmosphere. Itsrocket engines operate under extremes of pressure, tem-perature, shock, and vibration (Gregersen, 2009). Asmall defect factor or damage might lead to accidentsand failure of the entire operation.

Space launch—failure incidents

An investigation showed that among the 4378 spacelaunches that were conducted worldwide, 390 launchesfailed (failure rate of 8.9%), with an associated loss orsignificant reduction of service life of 455 satellites(some launches included multiple payloads) (Chang,2005). A US Vanguard vehicle exploded 2 s after liftoffon 6 December 1957 due to the low fuel tank and lowinjector pressure that allowed the high-pressure cham-ber gas to enter the fuel system through the fuel-injectorhead. A fire started in the fuel injector, destroying theinjector and causing a complete loss of thrust immedi-ately following liftoff. The US Saturn V had a singlefailure in the Apollo 6 mission on 4 April 1968, when

Figure 1. A Delta II launch vehicle with various structuralcomponents.Source: Pelt (2006).

Lee and Dhital 5

the third-stage engine failed to restart because of fuelinjector burn through. The Space TransportationSystem also suffered a single launch failure on 28January 1986, when the Challenger, carrying a seven-member crew, exploded 73 s into flight. At low tem-perature, the rubber O-rings in the motor case joint losttheir resiliency, and the combustion flame leakedthrough the O-rings and case joint, causing the vehicleto explode. Newly developed US commercial launchsystems, including Delta III, Conestoga, Athena, andPegasus, suffered launch failures during their earlydevelopmental flights in a repeat of the Vanguard,Juno, Thor, and Atlas failures in the late 1950s andearly 1960s (Chang, 2005). One Commonwealth ofIndependent States (CIS)/USSR space launch failureinvolved an SL-12 Proton vehicle carrying a Mars-96spacecraft on 16 November 1996. The second burn ofthe Proton’s fourth stage did not take place, and thespacecraft did not reach the interplanetary trajectory.Another failure involved the new Ariane-5 vehicle on 4June 1996. It veered off its flight path and exploded atan altitude of 3700 m only 40 s after liftoff. The failurewas attributed to errors in the design and testing of theflight software. Chinese space launch had a cata-strophic failure during the launch of the CZ-3B vehiclecarrying a commercial satellite, Intelsat 708, on 14February 1996. The vehicle and its payload hit theground and exploded in an inhabited area near thelaunch site 22 s after liftoff. The cause of failure wastraced to the CZ-3B’s guidance and control subsystem.A gold–aluminum solder joint in the output of one ofthe gyro servo loops failed, cutting electrical currentoutput from the power module and causing the inertialreference platform of the vehicle’s guidance and controlsystem to slope. Japanese liquid-propellant rockets (H-II) suffered two launch failures during 1998 and 1999.Japan’s other seven launch failures (including fourLambda-4S failures during the period 1966–1969)involved solid-propellant rockets (Chang, 2005). Theseare few of the examples among many such space launchfailure incidents.

Causes/sources of failure

Available launch-failure data reveal much about thepatterns in the possible causes of failure. Some failurecauses fall into the category of human error, such aspoor workmanship, judgment, or launch-managementdecisions, and some failures are the result of defectiveparts. Failure can occur in any phase of launch vehicledevelopment: inadequate designs and component tests;improper handling in manufacturing and repair pro-cesses; and insufficient prelaunch checks. Many pastfailures could have been prevented if rigorousreliability-enhancement measures had been taken(Chang, 2005). Fuel leaks (resulting from weldingdefects, tank, and feedline damage, etc.), payload

separation failures (from incorrect wiring, defectiveswitches, etc.), engine failure (a result of insufficientbrazing in the combustion chamber, from combustioninstability, hydrogen injector valve leakage, cloggedfuel lines, etc.), and loss of vehicle control (lightning,damaged wires that caused shorts, and control-systemdesign deficiencies), short circuits, engine thrust loss,software design errors that resulted in guidance systemfailure, wind shear, and residual propellants are amongthe causes that were revealed by previous failure analy-ses. Launch vehicle failure is usually attributed to prob-lems associated with a subsystem, such as propulsion,avionics, separation/staging, electrical component, orstructures. In some cases, failure is ascribed to prob-lems in another area altogether (e.g. launchpad, groundpower umbilical, ground flight control, and lightningstrike) or to unknown causes (usually when subsystemfailure information is not available) (Chang, 2005).

Flaws and damages in the launch vehiclecomponents

Various flaws or damages might occur on the launchvehicle structures and components during the manufac-turing/operational phase. This review study lists possi-ble patterns of damage that can be considered as thehot spot locations. It would be beneficial to determineand develop reliable nondestructive evaluation (NDE)and structural health monitoring techniques.

Payload fairing

The payload (spacecraft or satellite) is almost alwaysattached to the top of the vehicle. Payload fairingencloses and protects the payload during ground opera-tions and against atmospheric impact during launchvehicle ascent, against aerodynamic heating, andagainst harsh acoustic environment due to aerody-namic noise and jet noise (Gregersen, 2009; Gupta etal., 2007). It also forms the aerodynamic shape of thevehicle (Gupta et al., 2007), and in some cases, it alsoattenuates the acoustic loads from the main enginesduring launch (LM-3A Series launch vehicle user’smanual, 2011). A general payload fairing structuralconfiguration is shown in Figure 2. Once the launchvehicle is beyond the densest part of the atmosphere,the fairing is shed off (Gregersen, 2009).

The payload fairing experiences compression, ten-sion, shear as well as internal pressure. The most signif-icant loads on the fairing during operation are thecompression and bending due to the aerodynamic pres-sure and inertia produced by acceleration. Drag pro-duces a fairly uniform compressive stress throughoutthe shell, whereas bending loads produce tension, com-pression, and shear stress to different parts of the shell(Wegner et al., 2002). Moreover, various stresses are

6 Journal of Intelligent Material Systems and Structures 24(1)

induced from acoustic loads, thermal loads, jettisonloads, and ground handling loads. It should thereforemeet strength, buckling, flutter, thermal, and acousticrequirements and should also meet the design condi-tions without requiring significant modifications toexisting launchpad integration facilities. The fairingmust withstand the flight loads without experiencingany material failure, structural instability, or deflec-tions larger than allowed by the dynamic envelope ofthe fairing (Wegner et al., 2002). Fairing can be madefrom composites or metals in various structural designsand configurations as presented in the following.

Sandwich structure. Aluminum honeycomb core sand-wich structures were used to construct the payload fair-ing in Taurus II, Atlas V, and Ariane 5 launch vehicles.These fairings were of typical bi-conic design, made ofgraphite/epoxy face sheets with an aluminum honey-comb core. In this construction method, low mass andhigh stiffness can be combined. However, compositematerials are not isotropic (making damage growth pre-diction a complex process) and lack ductility (eliminat-ing yielding, which is an indicator of imminent failure).Also, the initiation and growth of material-level dam-age and the failure modes of composite structure arenot as well understood and cannot be predicted accu-rately (Harris et al., 2003). Fluid ingress and corrosionin the honeycomb region have also been reported (Hanet al., 1999). A well-known problem of honeycombpanels is the water trapping in the hexagonal cells of thehoneycomb, which corrodes the core material and soft-ens the composite face sheets (Huybrechts et al., 1999;Wegner et al., 2002). During the process of productionand service, defects such as delaminations and debondcan also appear. Delaminations can originate fromimpact damage to the composite skins. Because thesedefects are usually hidden in the structure, the evidence

of their existence will not appear before the skin–honeycomb core plates crack or break down (He et al.,2008). It is therefore very important to carry out non-destructive testing (NDT) of skin–honeycomb coresandwich structures before and after the assembly. Thedebonding between the inner skin and the honeycombcore is still one of the challenges in the inspection of anassembled payload fairing. Therefore, the inner skinand their interface should be inspected at least beforethe assembly.

Grid-stiffened structures. Composite grid-stiffened struc-tures are fabricated using a continuous fiber, organiccomposite material. A carbon fiber composite fairingconsists of a laminate skin co-cured to a reinforcingadvanced grid-stiffened (AGS) rib structure, as shownin Figure 3 (Biskner and Higgins, 2005; Higgins et al.,2004). These structures are characterized by a shellstructure (or skin) supported by a lattice pattern (orgrid) of stiffeners (Huybrechts et al., 1999; Van andWegner, 2002). With this technology, stronger, lighter,and low-cost structures, which can carry more payloadsto orbit, can be constructed (Kleiman, 2007). In con-trast to honeycomb sandwich panels, grid-stiffenedpanels do not trap water since the panel has a skin ononly one side, and moreover, can be manufacturedusing an almost entirely automated process. The auto-mated manufacturing has resulted in almost 20% costsavings, when compared to the construction cost ofhoneycomb sandwich payload fairings (Huybrechts etal., 1999; Van and Wegner, 2002; Wegner et al., 2002).Aluminum isogrid was the earliest precursor of themodern AGS structure. Isogrid refers to the triangulargrid pattern formed by stiffeners on the inboard side ofthe fairing structure. Such an isogrid structure, which ismachined from a single piece of aluminum stock, con-sists of a skin with stiffeners that form equilateral trian-gles. Despite being developed several decades ago, thisstructure is still used, as it was for the Titan, Atlas, andDelta launch vehicle shrouds and interstages. While avery proven and reliable structure, the aluminum iso-grid is heavy and expensive by today’s standards.Hence, grid-stiffened structures with shells supportedby a grid lattice of stiffeners have been researched formany decades as a possible replacement to monocoque,skin-stringer, aluminum isogrid constructions, andhoneycomb sandwich structures (Huybrechts et al.,1999). AGS structures take advantage of the high spe-cific stiffness and strength of carbon fiber materials byorienting the fibers along the rib direction. Thus, thesestructures are highly competitive with optimized honey-comb sandwich structures in terms of structural effi-ciency, that is, stiffness-to-weight and strength-to-weight ratios (Wegner et al., 2002).

However, these structures are also susceptible to var-ious structural failures. The three most common failure

Figure 2. A typical payload fairing structural configuration.Source: LM-3A Series launch vehicle user’s manual (2011).

Lee and Dhital 7

modes of AGS payload fairing structures are rib crip-pling, skin pocket buckling, and global Euler buckling.The majority of the applied load is carried by the ribs,and the skin simply stabilizes the ribs against buckling.However, since the skin has very little stiffness, itspocket buckles very easily. In post-buckling analysis,the structure is assumed to fail when the model becomesnumerically unstable and can no longer support addi-tional load (Wegner et al., 2002). A secondary failurecould occur when the interface stress exceeds thestrength of the epoxy bond between the rib and the skinis exceeded. Panel failure occurs from a fracture or dis-bond in this rib–skin interface. At this point, when therib separates from the skin, it starts to buckle since it nolonger is supported laterally against column buckling.The rib then buckles and fails (Wegner et al., 2002).Increasing skin thickness to control skin bucklingreduces the strains between the skin and ribs. Skinbuckling occurs in patterns that tend to peel the skinfrom the ribs. Once a rib disbonds, the structural integ-rity of the structure is compromised (Higgins et al.,2004).

Failure criteria thus include maximum strain, globalbuckling, joint failure resulting from shear, joint failureresulting from skin pull-off (tension), and joint failureresulting from skin peeling or bending. While the firsttwo criteria are common to all composite structures, thelast three are specific to grid-stiffened panels (Higgins etal., 2004). The other failure modes are shear joint fail-ure, which occurs when the shear load is transferred tothe skin through the adhesive layer; pull-off joint fail-ure, which occurs when the tensile stress exceeds thetensile strength of the adhesive layer between the riband skin; and peel-off joint failure, which is the criticalcriterion for many aerospace structures that uses thinskins. Bending moment causes the skin to peel-off from

the rib. Skin buckling causes a sharp increase of thebending moment. Skin pocket buckling induces a bend-ing load at the rib/skin joint (Higgins et al., 2004) asshown in Figure 4. The bond line between rib and skinis a layer of resin with resin pockets. These resin pock-ets are where the maximum skin bending momentoccurs and where failure is believed to initiate. The jointfailure due to skin peel-off was identified as the criticalfailure mode by one of the research studies of theMinotaur launch vehicle fairing.

Rocket propellant tanks/fuel tanks

Rocket propellant tanks are pressure vessels where pro-pellants are stored prior to use. They are often con-structed of materials such as aluminum alloys, steels, orcarbon fiber wound. The fuel and oxidizer tanks areusually of very lightweight construction, as they oper-ate at low pressures (Gregersen, 2009). Propulsion sys-tems, in particular tanks, are exposed to extremelyreactive and aggressive fluids, such as fuels, oxidizers,and cleaning agents that can promote stress corrosioncracks (SCCs). SCC is a failure phenomenon thatoccurs in engineering metallic materials by slow envir-onmentally induced crack propagation. SCC is a veryinsidious phenomenon since cracks can initiate andpropagate with little outside evidence (Ghidini et al.,2009). SCC of stainless steel bellows of a satellitelaunch vehicle propellant tank assembly was investi-gated (Jha et al., 2003). Bellows made of austeniticstainless steel (American Iron and Steel Institute (AISI)304 grade) are used as a conduit for liquid fuel and oxi-dizer in the propellant tanks of satellite launch vehicles.Austenitic stainless steel of type AISI 304 is prone toSCC in chloride environments (Jha et al., 2002). Thesebellows encountered frequent leakage problems during

Figure 3. Composite grid-stiffened payload fairing of Minotaur launch vehicle.Source: Higgins et al. (2004) and Biskner and Higgins (2005).

8 Journal of Intelligent Material Systems and Structures 24(1)

storage and use, as shown in Figure 5 (Jha et al., 2003).Cracks were found to originate from weld fusion lines.Macroscopic and microscopic observation revealed pitsat various locations on a weld line. Cracks were foundto originate from these pits. The formation of pits andthe crack initiation and propagation from these pits aretypical examples of SCC as shown in Figure 6 (Jhaet al., 2003). Furthermore, the rigidity of the ring atone end and the flexibility of plies on the other end leadto stress distribution during welding of bellows. Thepresence of cracks on the ring surface of the bellow,

very near to the weld fusion line, is attributed to thiscomplex stress that is developed due to the nonunifor-mity of the weld pool (Jha et al., 2006). AISI 304 gradesteel is also used as the material for the bolts used inthe flange joints of fuel and oxidizer pipelines in spacelaunch vehicles. Cracks were found in the shank at thehead–shank junction as well as in the threaded regionof these bolts, which resulted in bolt failure (Jha et al.,2002). There is always a possibility for missing thesefeatures during visual inspection, ultimately affectingfunctional performance later.

Figure 4. Skin pocket buckling induces a bending load at the rib/skin joint. (a) Skin bending and resulting rib and skin failuresurfaces. (b) Photomicrograph of rib/skin joint. The bond line between rib and skin is a layer of resin with resin pockets in the radii.These resin pockets are where the maximum skin bending moment occurs and where failure is believed to initiate. (c) Globalbuckling of forward cone. (d) Rib buckle and joint failure.Source: Higgins et al. (2004).

Lee and Dhital 9

Another research study investigated the cracking ofthe alloy nozzle weld of a propellant storage tank asshown in Figure 7 (Jha et al., 2009). AFNOR 7020(Al–4.5Zn–1.5Mg) is an aluminum alloy that is used tofabricate propellant tanks for satellite launch vehicles.Shrink fit nozzles are the nozzle openings for the liquidconduit. In one such tank, a crack developed at thenozzle weld region during storage. A metallurgicalanalysis revealed that the crack had developed due toSCC, which was promoted by the formation of aliquated region during welding operations. Aluminumalloys are also known to be susceptible to liquationcracking. Liquation cracking in aluminum alloy weldshas received much attention recently (Jha et al., 2009).

Lightweight structures and materials are consideredto meet the increased demand of performance of futurelaunch vehicles. Al–Li alloys such as 2195 are candidatematerials for reducing the structural weight of the cryo-genic tanks on launch vehicles due to their low density,high strength, and fatigue crack growth resistance(Moreira et al., 2010). However, the addition of Li toAl alloys presents problems, including low ductility andfracture toughness, delamination, and poor SCC resis-tance, which leads to fatigue crack susceptibility(Moreira et al., 2010). Another area for significantweight reduction is the replacement of metallic cryo-genic fuel tanks with composite tanks (Gates et al.,2006; Nair and Roy, 2006). Composite materials, which

Figure 5. (a) A stainless steel cut open bellow and (b) schematic diagram of joint configuration and location of leakage.Source: Jha et al. (2003).

Figure 6. (a) Macrographs showing pit, crack initiation from the pit, and crack extension toward the ring and also along the fusionline and (b) corrosion attack along the weld fusion line.Source: Jha et al. (2003).

10 Journal of Intelligent Material Systems and Structures 24(1)

have high specific strength and stiffness, and low coeffi-cients of thermal expansion (CTEs), have recently beeninvestigated for their applicability to propellant tanks.Under cryogenic environment, composite materialswith different CTEs between its reinforcing fibers andmatrix cause extreme thermal stress, which leads even-tually to microcracks in the polymer matrix (Kang etal., 2008; Nair and Roy, 2006). As the microcrack den-sity increases, they can propagate in the thickness direc-tion and emerge as a mechanism for leakage as well asmechanical degradation of the tank structure (Kang etal., 2008).

A promising structural concept based on polymericmatrix composite (PMC) materials is to use a sandwichconstruction with the PMC as the facesheet materialand lightweight polymeric honeycomb materials as thecore. Composite tanks based on this concept are foundto be used in the DC-XA experimental vehicle, NASAX-33 reusable flight demonstration vehicle, andSR-XM ELV. However, sandwich construction has anumber of potential problems because the multipleinterfaces may become a source for failure initiationand growth (Gates et al., 2006). One possible failuremechanism associated with a sandwich construction insuch a demanding environment is the debonding of thefacesheet. This mechanism can occur due to the highgas pressure inside the core material as a result ofhydrogen leakage through the facesheet. This mechan-ism is facilitated by the occurrence of cryopumping.The most likely initial failure mode due to cryopumpingis facesheet-to-core debonding, which can lead to crackgrowth and a total separation of the facesheet. Thisfailure mode occurred in NASA X-33, where the outerfacesheet and core separated from the inner facesheetalong a part of the tank wall. The pressure, coupledwith bond-line defects, likely caused the debond failure.For sandwich materials, there are two types of debond-ing near the skin–core interface: interface and

subinterface. In interface debonding, interface cracksgrow between the skin and core, and in subinterfacedebonding, subinterface cracks form when an initialinterface crack kinks into the core and then propagatesparallel to the skin within the core (Gates et al., 2006).

A fundamental study was carried out on the delami-nations in carbon fiber–reinforced polymer (CFRP)composite propellant tanks (Mizutani and Shimoda,2004). The lined filament wound (FW) CFRP tank willbe used for future space vehicles such as the GX-launchvehicle (Japan). However, the outer CFRP can delami-nate from the inner aluminum liner, which can lead tobuckling of the inner liner of tanks (Mizutani andShimoda, 2004). Severe thermal stress at the cryogenictemperature of liquid hydrogen can induce matrixcracks at relatively low loads (Kumazawa et al., 2006).The continuous chain of connected matrix crackswould cause a gas leakage, and a leak through dam-aged laminates would be significantly greater than aleak by diffusion through undamaged laminates. Gaspermeability through damaged CFRP composite lami-nates is influenced mainly by the densities and openingdisplacements of microcracks, which act as leakagepaths. Leakage of propellant is especially critical in thecryogenic composite tanks. In the case of cryogenicpropellant storage of liquefied hydrogen or oxygen, thedefects in the tanks that are induced during their manu-facturing and operation will lead to propellant leakageand may become the cause of structural fracture suchas the tank failure of the X-33 test vehicle through theaccumulation of gases in the honeycomb core(Kumazawa et al., 2006). One of the research studiesdescribes short cracks (also called stitch cracks) in mul-tidirectional composite laminates used for cryogenicpropellant tanks. It has been observed that the patternof cracks in multidirectional laminates is quite differentfrom that in cross-ply laminates (Xu and Sankar,2008). Impact also tends to create several other damage

Figure 7. Optical photomicrographs showing (a) cracking very near the fusion line and (b) thick film along the grain boundaries.Source: Jha et al. (2009).

Lee and Dhital 11

factors in a laminate, as shown in Figure 8 (Will et al.,2002). Fatigue failure is also one of the flaw phenom-ena in a multidirectional laminate, and during fatiguefailures, various other damages can develop, whichinclude matrix cracks, layer delaminations, interfacefailure, and fiber fracture (Vassilopoulos and Keller,2011).

In some cases, a Type 3 composite tank is generallypreferred to a Type 4 tank. A Type 3 is a compositetank that is lined with metal to maintain permeability,and a Type 4 is a tank that is fabricated only withcomposite materials (Kang et al., 2008). Hoop andhelical layers of composite are wound onto the metalliner in the Type 3 tank structure. In this filamentwinding process, acetone is used to lower the viscosityof the resin. However, delaminations occur in thecomposite laminate of the cryotank, as shown inFigure 9, mainly due to the voids caused by the eva-poration of acetone during the fabrication procedure(Kang et al., 2008). Microscopy showed that someinterfaces between the hoop and helical layers aredetached locally (delamination), as shown in Figure9. Problem during the fabrication process also leadsto many defects near the interface between hoop and

helical layers, such as voids and microcracks, asshown in Figure 10.

Payload adapter

A payload adapter functions as the interface betweenthe mounted payload and the space launcher body. Thepayload adapter includes mechanisms that will free thepayload from the launcher when the space vehicle hasreached orbit (Pelt, 2006).

Composite adapter for shared payloads (CASPAR). CASPARis a multi-payload adapter (MPA) developed withwhole-spacecraft vibration isolation for launch vehi-cles, such as the Atlas V/Delta IV and Minotaur IV.With an optimized available payload envelope, thisstructure was designed to carry multiple payloads(Gies, 2009; Maly et al., 2005; Sanford et al., 2009), asshown in Figure 11. In other words, it was designed tosupport co-manifested primary payloads, allowing onerocket launch to support two satellite launches. It is acomposite design of solid graphite/epoxy laminate shelldeveloped to have lightweight, integrated vibration

Figure 9. (a) Composite tank from filament winding process and (b) microscopy of the cross section of the cylinder part of thecryotank.Source: Kang et al. (2008).

Figure 8. (a) Illustration of impact damage in a laminate and (b) illustration of the propagation of delamination in a laminate.Source: Will et al. (2002).

12 Journal of Intelligent Material Systems and Structures 24(1)

isolation, low shock, and modularity. The reinforce-ment pad-up around the access door perimeter consistsof doubler plies on the outside of the MPA structure

and a bonded aluminum doubler on the inside surface(Maly et al., 2005). CASPAR accommodates two 1500-lb satellites, or up to four ELV secondary payload

Figure 10. (a) Microscopy of the interface between hoop and helical layers of the composite/aluminum ring before and (b) afterliquid nitrogen (LN2) immersion test.Source: Kang et al. (2008).

Figure 11. (a) CASPAR used in Minotaur IV launch vehicle, (b) CASPAR finite element model, and (c) CASPAR with co-manifestedprimary spacecraft in fairing of Minotaur IV.CASPAR: Composite adapter for shared payloads.

Source: Maly et al. (2005).

Lee and Dhital 13

adapter (ESPA)-class satellites when used with one ortwo flat-plate Adapters (FPAs) (CSA Engineering,Online access). CASPAR is approximately 60-pliesthick, made of industry-standard carbon/epoxy prepregcomposite material. The major diameter tapers to theall-composite flange at both ends (Gies, 2009).

Experimental and analytical failure tests were con-ducted to understand composite failure in CASPAR(Biskner et al., 2009). The bond between the accessdoor doublers and the structural shells was predicted tofail under compressive loading before the failure of theoverall structure, and this bond failure will in turncause the unsupported door cutouts to buckle, resultingin the overall structural failure. Composite skin wrink-ling could occur due to tension of the CASPAR shell(Biskner et al., 2009). Local matrix failure, local fiberfailure, and ultimate failure, which reduce stiffness, arecommon in composites (Nelson and Welsh, 2009). Alsothe prepreg tape can wrinkle in the region of the for-ward and the aft radii. Wrinkling creates fiber wavi-ness, which could have a significant effect on thestrength or stiffness of CASPAR (Nelson and Welsh,2009). Figure 12 shows the matrix and fiber failures inthe CASPAR structure.

Interstage adapter. Interstage adapter (ISA) provides thestructural connection between the booster stage andthe upper stage and was used in Atlas V, as shown inFigure 13. Both the forward and aft flanges are metallicand are attached to the composite cone via fasteners.Two thin composite face sheets surround an aluminumhoneycomb core in a sandwich structure to compose

the shell of the ISA. The composite used in the ISA isan industry-standard carbon/epoxy cocured on thecore. The ISA assembly models include a compositeconic adapter, a forward and aft attach ring, and a loadhead. The body of the ISA is a composite sandwichstructure containing thin carbon/epoxy composite facesheets with an aluminum honeycomb core. Matrix andfiber failures in ISA structures have been reported, asshown in Figure 14 (Nelson and Welsh, 2009).

TPS

TPS is the barrier that protects the external structuresof launch vehicles from their rough encounter withatmospheric layers. It also protects these structuresfrom the heat and cold of space while in orbit. It con-sists of various materials that are applied to the outerstructural skin of the launch vehicle to maintain theskin within acceptable temperatures. Thermal protec-tion also acts as mechanical protection during animpact and reveals possible structural damage locations(Bouvet et al., 2004). All space launch vehicles consistprimarily of large tanks that hold propellants. Thesecryogenic propellant tanks and feed lines must be ther-mally insulated to prevent or minimize air condensa-tion or ice formation as well as to provide sufficientisolation of the components from the ambient environ-ment (Fesmire et al., 2011). The system covers most ofthe rocket tank’s external surface to help keep theliquid hydrogen and oxygen propellants inside at cryo-genic temperatures. As a whole, it maintains appropri-ate propellant temperatures, protects the skin from

Figure 12. Side view of CASPAR showing an envelope failure quilt plot. Blue elements indicate no failure, green elements indicatematrix failure, and red elements indicate fiber failure.CASPAR: Composite Adapter for Shared Payloads.

Source: Gies (2009).

14 Journal of Intelligent Material Systems and Structures 24(1)

Figure 13. ISA in Atlas V launch vehicle.ISA: Interstage Adapter.

Source: Nelson and Welsh (2009).

Figure 14. (a) Failures in the ISA structure. Failure initiated at the upper corners of the modified door and immediately propagatedaround the circumference of the article. Both the inner and outer face sheets experienced similar failures. (b) Analytical predictionsof failure propagation. Green symbolizes matrix failure while red indicates fiber rupture.ISA: Interstage Adapter.

Source: Nelson and Welsh (2009).

Lee and Dhital 15

aerodynamic heat, and minimizes ice formation. Thethermal insulation system is designed to meet threeinterrelated requirements: safety, control, and flightperformance. Safety aspects include the prevention ofexcessive ice formation or liquid air condensation,which can present a debris hazard or an enhancedflammability problem. Control during prelaunch load-ing operations is of course vital for successful cooldown, tanking, stabilization, and replenishment withinthe time constraints of the mission. The insulation sys-tem provides the added benefit of reducing the continu-ous boil-off losses of the cryogenic propellants. Flightperformance requirements include thermal protectionof the vehicle tanks from radiant heating or aerody-namic effects (Fesmire et al., 2011).

Low-density polyurethane foam has been an impor-tant insulation material for space launch vehicles forseveral decades (Workman et al., 2007). The Saturnlaunch vehicle used polyurethane foam as its exteriorfuel tank insulator, and it was at this time that theexperience with this unique material began. In theJapanese H-1 launch vehicle, the fuels were containedin the 2219 Al alloy tank insulated with sprayed poly-urethane foam. In the Soyuz space launch vehicle, thepayload fairing consisted of a two-half-shell CFRPsandwich structure with an aluminum honeycomb core.The total thickness of the fairing was approximately 25mm. A 20-mm-thick thermal cover made of polyur-ethane foam with a protective liner was applied to theinternal surface of the cylindrical part of the fairing.The potential for damage from the foam’s breakingaway from the NASA Space Shuttle ET was not rea-lized until the impacts from the foam that broke awayon the Columbia orbiter vehicle caused damage to thewing’s leading edge. Voids occurred in such foams(Workman et al., 2007). Other cryogenic tank insula-tion materials that were used in the past included rigidfoams and cork (Fesmire et al., 2011). Cork thermal

shield has been used on composite panels (Bouvet etal., 2004).

Spray on foam insulator (SOFI) is also one of thewidely used thermal protection material of cryogenicinsulation. SOFI is closed-cell foam, and by its nature,is mostly air (Bridges, 2005). It is durable and has lowthermal conductivity and density (Barrios and Sciver,2010). In the ET of the Space Shuttle, it provided ther-mal insulation over the shuttle’s outer surface to pre-vent ice formation. Thermal cycling is an importantdesign parameter determining the mechanical behaviorof insulator and must be carefully considered to avoidadverse consequences such as cracking or delaminationon the tank wall (Fesmire et al., 2011). Since SOFI isvulnerable to weathering exposure and repeated ther-mal cycling, it is generally used for onetime-use vehicletanks.

Certain areas of the tanks have geometric disconti-nuities such as bolts, flanges, and fittings. When thefoam insulator is sprayed onto these areas, voids canbe created because of these geometric discontinuities,as shown in Figure 15 (Arakere et al., 2008; Workmanet al., 2007). Liquid nitrogen and oxygen (air) can con-dense into any of these voids. During liftoff, the outersurface of the foam is exposed to aerodynamic heating,which can raise the temperature of the liquid nitrogen,turning it into gas. The pressure change due to the for-mation of the gas can blow out the foam in pieces dur-ing liftoff (Arakere et al., 2008). Space ShuttleColumbia’s catastrophic failure was caused by a pieceof the ET insulating foam that separated from the ETand struck the leading edge of the left wing of the orbi-ter. Voids that form within the foam during the appli-cation of the foam on a tank can lead to cohesivefailure. The air entrapped within the voids can expandfrom the heat generated during ascent, increasing thepressure inside the foam, and ultimately cohesively fail-ing the foam between the void and the foam surface.Near or at the interface between the foam and the tankstructure where voids form, entrapped air will be lique-fied by the liquid hydrogen or liquid oxygen tempera-tures at the tank’s aluminum surface, so the potentialfor cohesive failure is exasperated. Small air voids mayalso cause separation of the SOFI from the ET duringa launch. Air voids in SOFI can significantly reducethe vehicle integrity (Case et al., 2005, 2006).

Two types of foam, namely, ‘‘acerage’’ material and‘‘hand sprayed’’ foam were used on the ET of theSaturn V rocket, which similar to the Space Shuttleused cryogenic fuel. The acerage foam posed small-scale cohesive failures, known as ‘‘popcorning’’ in theintertank region. It was determined that temperature(slow heating) and the differential between the foaminternal pressure and the external vacuum were thecauses of ‘‘popcorning.’’ Voids can be formed fromspraying the foam at too low a temperature. One of thetest studies outlined the following 11 defect categories:

Figure 15. Example of natural void formation. They lacksymmetry and straight lines or smooth surfaces.Source: Workman et al. (2007).

16 Journal of Intelligent Material Systems and Structures 24(1)

delamination from the primer, delamination from theadhesive, delamination along the knit line, delamina-tion from the substrate, delamination between therepair material and the adhesive, the presence of con-centrated material ejected by gun spit, formation ofelongated cells, porosity, crack, rollover, and voids(Weiser et al., 2004).

Damage evaluation necessity andconventional techniques

Damage evaluation necessity

Detection of damage and prediction of strength lossfrom expected damage are critical design considerationsand are an important issues for the launch operation ofa space launch vehicle. Each failure puts the spacelaunch operation ‘‘on hold’’ until the vehicle canresume service. A typical cost estimate of a failure forAriane 4 was US$466 million (total launcher systemcost)+US$79 million (total customer costs). The‘‘downtime’’ associated with a launch failure also variesconsiderably, for example, 3.7 months for Delta and8.1 months for Ariane. After the Challenger accident,Space Shuttle took 32 months to return to service(Parkinson, 1999).

Invisible flaws, such as impact damage or a partiallybonded joint, can lead to costly failures during a launchor in orbit, where components generally cannot befixed. Detection of flaws is therefore essential.Moreover, flaws need to be discovered nondestruc-tively, that is, the process used to test for them mustnot damage the potentially unflawed component beingtested. The process of flaw detection and characteriza-tion is known as NDE, and it is an important part ofensuring the flightworthiness of launch vehicles andspacecrafts. It begins with the inspection of raw materi-als, both to check material characteristics and to screenfor incipient flaws. It continues through manufacturingand assembly as a means of process control and qualityassurance. It concludes with a final check of compo-nents for damage incurred during shipping, storage,and handling (Johnson and Esquivel, 2006). An earlydetection of faults can prevent unnecessary systemshutdowns, breakdowns, and catastrophes involvinghuman fatality and material damage. The maintenancecost of industrial processes can therefore be reduced,and product quality can be enhanced (Wu, 2004).

Conventional techniques

Various types of NDE can be applied, depending onthe component and material. Critical inspection para-meters include the geometry and constitution of theinspected part and the type of defect to be detected. Ifsurface cracks are the primary defect of concern, apenetrating dye can be applied to make the cracks more

visible. To check for cracks just below the surface, aneddy current probe can be used to look for the changesin conductivity associated with subsurface flaws (assum-ing the material is conductive). To look for even deepercracks, in a weld line, for example, the best approachmight be to direct ultrasonic shear waves into the weldand look for characteristic echoes. In the aerospaceindustry, two of the most widely used techniques fordetecting volumetric (subsurface) flaws are radiographicand ultrasonic testing. In radiographic testing, X-ray,gamma ray, or neutron is directed at a part and detectedafter it passes through part. Flaws give rise to shadowsor bright spots in the detection field recorded on film ora detector array. Typical applications of radiographictesting include detection of voids in solid propellants,casting ingots, and adhesives; assembly verification;detection of core anomalies in honeycomb panels; andinspection of welds (Johnson and Esquivel, 2006). Laserultrasonic testing capable of wave propagation imagingand damage visualization can also be applied to compo-site and metallic components or full-scale space launch-ers during assembly (Chia et al., 2011).

Other techniques like thermography, shearography,and acoustic emission monitoring are also available. Inthermographic testing, heat is passed through theinspected part, and subsurface defects, if present, willimpede or enhance the flow of heat, resulting in loca-lized hot or cold regions on the surface. Applicationsinclude detecting bond-line flaws beneath compositeface sheets in honeycomb structures and detecting heatleaks. Shearography is a field-applicable holographictechnique that can be used to identify the regions ofvery fine relative displacement or strain. A special cam-era takes pictures of the reflection of diffused laser lightfrom the test article surface before and after it is sub-jected to a mechanical stress, such as slight pressuriza-tion. These pictures are processed to yield an imageconsisting of fringe patterns. Regions of concentratedsurface strain under load result in a higher density offringes. Fringe pattern anomalies are indicative of sub-surface flaws. This method could be used to detectimpact damage in the spherical composite overwrappedpressurization vessels used to contain compressed gasesin a number of spacecrafts. Acoustic emission monitor-ing can be performed during hardware proof tests.Small microphones, sensitive to sound at frequenciesbeyond the audible range, are affixed at strategic loca-tions on the test article. The acoustic transducers aremonitored for sound emissions emanating from thehardware throughout the loading cycle. Cracks andother flaws that grow during the proof test make noisethat is picked up by these sensitive microphones, andthe location of these flaws can be determined throughtriangulation (Johnson and Esquivel, 2006). Acousto-ultrasonics are similar to the acoustic emission tech-nique, in which a structure is loaded and the structureis monitored by transducers. This technique was used

Lee and Dhital 17

for events such as fiber breaking, matrix cracking, anddelaminations, which produce ultrasonic signals asstrain energy associated with the events is released(Paris, 1994). Eddy current inspection is based on gen-erating or inducing electrical currents in a workpieceand understanding how they respond to the presence offlaws or material changes such as subsurface cracks,voids, impurities, and corrosion. The technique wasused in one of the research studies of the cracked flightnozzle on a Delta rocket’s graphite epoxy motor, whoseliner was detected to be out of tolerance (Johnson andEsquivel, 2006).

However, these conventional techniques pose somelimitations as well. NDE of large structures with closedsurfaces such as filled pressure vessels is difficult.Ultrasonic techniques such as C-scan require large fix-tures and usually access to both sides of a structure.The problem with acoustic emission is that there is noclear spatial location for ‘‘events,’’ and it is difficult tocorrelate specific types of damage in complex structureswith specific acoustic emission events. A major diffi-culty with acousto-ultrasonic is the need to couple theacoustic signal to the structure either by immersion of

the structure in a fluid or by use of oil, grease, or othercoupling media. The presence of surface roughness andcontamination can change reflectivity and emissivity,which changes the energy transfer to the structure(Paris, 1994). NDE techniques including the ultrasonictechnique, X-ray technique, and the eddy current tech-nique can only be applied on disassembled components(Qing et al., 2006). Hence, several researches are beingconducted to further study and improve the conven-tional NDE techniques and to develop technologies tosuit aerospace requirements for zero-defect spacelaunch structures.

Future works

Optical NDT has gained more and more attention inrecent years, mainly because of its characteristics withhigher detection accuracy and sensitivity and ease ofsignal multiplexing and resistance to electromagneticinterference. Some of the main optical NDT technolo-gies are fiber optic technology, terahertz technology,and fully noncontact scanning excitation lasertechnology.

Fiber optics feature easy integration and embedding,and terahertz technology opens a new direction of inter-nal NDT because of its excellent penetration capabilityto most of nonmetallic materials (Zhu et al., 2011).Fiber optic sensing technique also shows potential, andone of the research studies mentions it for applicationin next-generation launch vehicles. This technique coulddevelop multiparameter and high-temperature multi-plexed optical fiber sensors for metallic and compositestructures. In composite cryotank of space launch vehi-cle, high spatial densities of fiber Bragg grating sensorarrays are bonded on the exterior of the compositetank. To enable integrated vehicle health managementfor the next-generation launch technology cryotank, alightweight multisensor technology is needed as a viablesubstitute for conventional strain gages that are notpractical for such application Glass (2008).

Terahertz sensing and imaging provides sufficientphase and amplitude contrast for the study of the spec-tral properties of targets in the terahertz domain(Zhong, 2006). Noncontact imaging in combinationwith spectroscopic capabilities can provide additionalinformation of the sample under investigation. Criticalfeatures like thickness, shape, defects, or delaminationsare detectable. Most items consisting of dielectrics suchas plastics, glass fiber–reinforced plastics, or ceramicsare transparent for terahertz waves (Jonuscheit et al.,2010). One of the industrialized applications of tera-hertz imaging technique as of now is the NDT of theTPS in space launch vehicles (Zhong, 2006).

Another potential technique is shown in Figure 16Lee (2012). The scanning excitation laser coupled withnoncontact ultrasonic sensors such as laser Dopplervibrometers, air-coupled piezoelectric, and capacitance

Figure 16. Scanning excitation laser technology coupled withnoncontact laser ultrasonic sensor and contact or noncontactpiezoelectric ultrasonic sensor.Source: Lee (2012).

18 Journal of Intelligent Material Systems and Structures 24(1)

sensor has a good potential of NDE of full-scale spacelauncher structures in each assembly step and satisfiesspace structural manufacturers’ request of fully non-contact inspection.

Conclusion

Space launch vehicles have several systems and subsys-tems. Flaws during the manufacturing phase or damageinduced during assembly/ground handling in any of thecomponents increases the failure probability of thespace launch operation. Many of these failures willhave adverse consequences on safety, cost, life, and/orschedule. Hence, failure prevention is necessary up tothe point where the effort of failure prevention not out-weighs against the consequences of failure. Many ofthese failures can be prevented through proper design,testing, and operation. Since different materials areused for different components, the nature and types ofdamage also differ. Possible flaws and damages inmajor components like payload fairing, propellanttanks or fuel tanks, payload adapters, and TPSs wereinvestigated in detail for better understanding of therisks posed upon each major component.Understanding the unique risks posed upon each com-ponent might be useful for developing an efficient andreliable nondestructive system. Finally, the conven-tional and high potential techniques were reviewed toaddress the state-of-the art in the detection of suchflaws and damages.

Funding

This study was supported by the Leading Foreign ResearchInstitute Recruitment Program (2011-0030065) and BasicScience Research Program (2011-0010489) through theNational Research Foundation of Korea funded by theMinistry of Education, Science and Technology. In addition,this study was supported by the University CollaborationEnhancement Project of the Korea Aerospace ResearchInstitute.

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