aerospace Article
Large Scale Applications Using FBG Sensors: Determination of In-Flight Loads and Shape of a Composite Aircraft Wing Matthew J. Nicolas 1,† , Rani W. Sullivan 2,†, * and W. Lance Richards 3,† 1 2 3
* †
Department of Manufacturing Engineering, PACCAR Engine Company, Columbus, MS 39701, USA;
[email protected] Department of Aerospace Engineering, Mississippi State University, Mississippi State, MS 39762, USA NASA Engineering and Safety Center, NASA Langley Research Center, Hampton, VA 23681, USA;
[email protected] Correspondence:
[email protected]; Tel.: +1-662-325-7289 These authors contributed equally to this work.
Academic Editor: Rafic Ajaj Received: 16 May 2016; Accepted: 8 June 2016; Published: 23 June 2016
Abstract: Technological advances have enabled the development of a number of optical fiber sensing methods over the last few years. The most prevalent optical technique involves the use of fiber Bragg grating (FBG) sensors. These small, lightweight sensors have many attributes that enable their use for a number of measurement applications. Although much literature is available regarding the use of FBGs for laboratory level testing, few publications in the public domain exist of their use at the operational level. Therefore, this paper gives an overview of the implementation of FBG sensors for large scale structures and applications. For demonstration, a case study is presented in which FBGs were used to determine the deflected wing shape and the out-of-plane loads of a 5.5-m carbon-composite wing of an ultralight aerial vehicle. The in-plane strains from the 780 FBG sensors were used to obtain the out-of-plane loads as well as the wing shape at various load levels. The calculated out-of-plane displacements and loads were within 4.2% of the measured data. This study demonstrates a practical method in which direct measurements are used to obtain critical parameters from the high distribution of FBG sensors. This procedure can be used to obtain information for structural health monitoring applications to quantify healthy vs. unhealthy structures. Keywords: fiber Bragg grating; FBG; carbon composite wing; optical fiber strain measurement; flight loads; wing deflection; wing shape; structural health monitoring
1. Introduction Structural health monitoring (SHM) refers to the process of observing the structural and/or mechanical responses of a system over time for the purposes of damage detection and characterization. An effective SHM method can increase efficiency, reduce maintenance and inspection costs, and extend the service life of mechanical systems and structures. An overview of literature in the field of optical methods indicates that fiber Bragg grating (FBG) sensor technology is considered a promising and commonly used method for evaluating and monitoring structural system response and damage detection [1]. In the following sections, a general description of FBG technology is given, followed by an overview of large scale applications in which FBGs have been used. Additionally, a case study is presented to demonstrate a practical method in which direct measurements (strains from the high distribution of FBG sensors along the optical fibers) are used to obtain not only the wing shape, but
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also the out-of-plane or in-flight load distribution. The procedure can be used to obtain information for structural health monitoring applications. 2. Fiber Bragg Grating (FBG) Technology 2.1. FBG Features FBG technology quickly gained consideration as a measurement and monitoring tool due to its many attributes, such as the extremely small size and light weight of the sensors. A primary benefit of using FBG technology is the multiplexability of the optical sensors that enables the monitoring of a high-density strain distribution using a single fiber [2,3]. In applications that require a large number of sensors, significant cost savings per sensor is realized when compared to the installation cost of conventional foil strain gages. Unlike conventional sensors, FBGs are nonconductive, electrically passive, and immune to electromagnetic interference, enabling their use in environments subject to noise, corrosion, or high voltage [4]. These diminutive sensors can be directly integrated into polymer matrix composite (PMC) structures [5,6] or surface mounted [7] on the test object. FBGs, which have the capability to measure very large strains (>10,000 µm/m), can be used to monitor critical load bearing structures. Also, FBGs have long term stability, good corrosion resistance, and are not distance dependent [8]. In using FBG sensors, there are also issues that must be addressed. FBG sensors have high temperature dependence and at high temperatures, there can be distortion of the optical fibers due to the mismatch in the thermal expansion coefficients of the cladding, core and any coating. In mounting or attaching these sensors, the issue of ingress and egress of the optical fiber are major concerns due to the fragile nature of the fibers [9–11]. Additionally, the cost of optical interrogators, used for the measurement of FBGs, can be prohibitive for use on small aircraft. As such, research is ongoing for development of low-cost systems [12]. 2.2. Principle of FBG Sensors A fiber Bragg grating (FBG), first demonstrated by Hill [13], is a type of reflector etched in the core of an optical fiber that reflects particular wavelengths of light and transmits all others, enabling its use as an optical sensor. The center of the fiber is surrounded by a silica cladding with a slightly lower index of refraction than the core (4–9 µm diameter) to enable the formation of a wavelength inside the fiber. Each FBG is defined with a unique wavelength that changes with induced strains. The refractive planes in an FBG are called Bragg planes, which are produced when two opposing light sources interfere constructively. Stretching an FBG (due to mechanical or thermal strain) causes a shift in the sensor’s index of refraction. As illustrated in Figure 1, when a wavelength tunable laser propagates light through an FBG, a narrow bandwidth of the laser is reflected back causing interference, while the remaining is transmitted to the remaining FBGs. The FBG strain is obtained from the Bragg wavelength, λb , which is in the center of the reflected wavelength range, and can be expressed as [4,5,14]: λb “ 2 n Λb (1) In Equation (1), n is the effective refractive index of the fiber core, and Λb is the Bragg planes spacing. A shift in the reflected wavelength occurs when a Bragg grating sensor is elongated by an induced strain. Due to temperature change ∆T and strain ε in the fiber direction, the Bragg wavelength shift ∆λb can be expressed as [4]: ´ ¯ ∆ λb “ p1 ´ Pe q ε ` αs ` α f ∆T λb
(2)
where λb is the unstrained Bragg wavelength, Pe is the strain-optic coefficient, αs is the thermal expansion coefficient of the fiber, and α f is the thermal-optic coefficient. Since both temperature and
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and strain influence the reflected wavelength of temperature an FBG, temperature compensation must be strain influence the reflected wavelength of an FBG, compensation must be introduced introduced accurate strain measurements. for accurate for strain measurements.
Figure1.1.Principle Principle of fiber (FBG) showing light propagation, and Figure of fiber BraggBragg gratinggrating (FBG) showing light propagation, reflection andreflection transmission transmission through the core of the fiber. through the core of the fiber.
3. FBG Monitoring of Large Scale Structures 3. FBG Monitoring of Large Scale Structures FBG-based systems are enjoying wide usage in a number of industries and applications. The FBG-based systems are enjoying wide usage in a number of industries and applications. The monitoring of structural performance to obtain loads and detect damage is increasingly gaining monitoring of structural performance to obtain loads and detect damage is increasingly gaining importance for many applications, particularly in the aerospace industry. Typically, load importance for many applications, particularly in the aerospace industry. Typically, load monitoring is monitoring is accomplished by measuring local strains and damage detection involves the accomplished by measuring local strains and damage detection involves the monitoring of acoustic monitoring of acoustic signals [15]. signals [15]. The widest use of FBGs is for measurement of strain, temperature, and pressure. Primarily, the The widest use of FBGs is for measurement of strain, temperature, and pressure. Primarily, shift in the wavelength or the refractive index is measured by the optical system and the strain ε is the shift in the wavelength or the refractive index is measured by the optical system and the strain ε is computed by Equation (2). computed by Equation (2). A number of studies deal with large aerospace vehicles and their components and most work A number of studies deal with large aerospace vehicles and their components and most work involves the monitoring of strain data. The NASA Armstrong Flight Research Center (AFRC) has involves the monitoring of strain data. The NASA Armstrong Flight Research Center (AFRC) has developed an instrumentation system and analysis techniques that combine to make distributed developed an instrumentation system and analysis techniques that combine to make distributed structural measurements practical for lightweight vehicles. AFRC’s fiber optic strain sensing (FOSS) structural measurements practical for lightweight vehicles. AFRC’s fiber optic strain sensing (FOSS) technology enables a multitude of lightweight, distributed strain measurements [4,7,16]. Motivated technology enables a multitude of lightweight, distributed strain measurements [4,7,16]. Motivated by the failure of the unmanned aerial vehicle (UAV) Helios, Richards et al. [4,17] used the spatial by the failure of the unmanned aerial vehicle (UAV) Helios, Richards et al. [4,17] used the spatial resolution and equal spacing of FBGs to monitor the real-time in-flight wing shape of the Ikhana resolution and equal spacing of FBGs to monitor the real-time in-flight wing shape of the Ikhana and and Global Observer UAVs. Alvarenga et al. [18] also used FBGs on a lightweight UAV to Global Observer UAVs. Alvarenga et al. [18] also used FBGs on a lightweight UAV to determine the determine the real-time deformation shape. A number of studies performed ground testing on real-time deformation shape. A number of studies performed ground testing on aircraft and obtained aircraft and obtained measurements through FBGs mounted on the surface of the test structure. measurements through FBGs mounted on the surface of the test structure. Childers et al. [2] performed Childers et al. [2] performed static loads tests on an advanced composite transport wing to obtain static loads tests on an advanced composite transport wing to obtain strain measurements from strain measurements from thousands of FBGs and compared these to measurements obtained from thousands of FBGs and compared these to measurements obtained from electrical foil gages. The high electrical foil gages. The high density of measurements revealed structural behavior that is typically density of measurements revealed structural behavior that is typically obtained from numerical models. obtained from numerical models. Kim et al. [12] used a low speed, low cost FBG interrogator to Kim et al. [12] used a low speed, low cost FBG interrogator to obtain temperature compensated strain obtain temperature compensated strain measurements of a full-scale small aircraft wing structure measurements of a full-scale small aircraft wing structure subjected to various temperatures and subjected to various temperatures and loading conditions. To develop a real-time UAV SHM loading conditions. To develop a real-time UAV SHM system, Gupta et al. [19] investigated a number system, Gupta et al. [19] investigated a number of issues regarding FBGs such as sensor integration, of issues regarding FBGs such as sensor integration, location, and embedment. They used actual location, and embedment. They used actual flight test data to monitor the vibration and loads flight test data to monitor the vibration and loads signature during flight conditions. As a part of signature during flight conditions. As a part of the European Smart Intelligent Aircraft Structures the European Smart Intelligent Aircraft Structures project, Ciminello et al. [20] developed an in-flight project, Ciminello et al. [20] developed an in-flight shape monitoring system that uses the chordshape monitoring system that uses the chord-wise strain distribution obtained from a network of FBG wise strain distribution obtained from a network of FBG sensors for an adaptive trailing edge sensors for an adaptive trailing edge device. device. FBG sensors can also be integrated into the test structure. Amano et al. [21] embedded FBG sensors can also be integrated into the test structure. Amano et al. [21] embedded multiplexed-fiber FBG sensors into an aerospace carbon/epoxy advanced grid structure to measure multiplexed-fiber FBG sensors into an aerospace carbon/epoxy advanced grid structure to measure mechanical strains of all ribs to evaluate damage under low-velocity impact loading. Dvorak et al. [22] mechanical strains of all ribs to evaluate damage under low-velocity impact loading. Dvorak et al. embedded FBG sensors into the carbon/epoxy spar caps and in the adhesive joints of a glass/epoxy [22] embedded FBG sensors into the carbon/epoxy spar caps and in the adhesive joints of a composite wing to monitor strains during a static loads test. Ruzek et al. [23] embedded FBGs in glass/epoxy composite wing to monitor strains during a static loads test. Ruzek et al. [23] embedded carbon-fiber reinforced plastic (CFRP) fuselage stiffened panels and performed compression after FBGs in carbon-fiber reinforced plastic (CFRP) fuselage stiffened panels and performed impact tests. The FBG-based system successfully acquired all buckling modes of the CFRP panels.
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compression after impact tests. The FBG-based system successfully acquired all buckling modes of Aerospace 2016, 3, 18 4 of 15 the CFRP panels. Other large scale applications include the work by Kim et al. [10] in which FBG sensors were embedded thescale adhesive layer joining thethe spar cap by to Kim the shear of which composite turbine Other in large applications include work et al. web [10] in FBG wind sensors were blades to obtain deflection measurements. FBG-based systems have been used in the structural embedded in the adhesive layer joining the spar cap to the shear web of composite wind turbine testing ofobtain composite overwrapped pressure vessels systems (COPVs). Banks al. [6] embedded FBGs to blades to deflection measurements. FBG-based have beenetused in the structural testing measure strain distributions in Kevlar COPVs during stress rupture tests. In-situ SHM was of composite overwrapped pressure vessels (COPVs). Banks et al. [6] embedded FBGs to measure performed on filament wound pressure tanks to water pressurization tests using strain distributions in Kevlar COPVs during stresssubjected rupture tests. In situ SHM was performed on embedded FBGs [24]. Mizutani et al. [25] conducted real-time strain measurements using FBG filament wound pressure tanks subjected to water pressurization tests using embedded FBGs [24]. sensors liquid hydrogen tank mounted on a reusable rocket In liquid other Mizutanion et aal.composite [25] conducted real-time strain measurements using FBG sensors on avehicle. composite transportation fields, such as the railway market, an FBG-based system has been developed and hydrogen tank mounted on a reusable rocket vehicle. In other transportation fields, such as the railway used on an commercial to monitor hard andand softused impacts on currentrailways collectors, enablinghard the market, FBG-basedrailways system has been developed on commercial to monitor detection of serious defects [26]. In obtaining the strain responses of steel wires in a column support and soft impacts on current collectors, enabling the detection of serious defects [26]. In obtaining of subway structure under ground motion, support Chen et of al.a[27] found that FBG sensors performed theastrain responses of steel wires in a column subway structure under ground motion, better than strain gages because the area is under large electromagnetic interferences. Wan [28] Chen et al. [27] found that FBG sensors performed better than strain gages because the areaetisal. under analyzed dynamic properties obtained from FBG data in the testing and monitoring of a large scale large electromagnetic interferences. Wan et al. [28] analyzed dynamic properties obtained from FBG truss bridge. data in the testing and monitoring of a large scale truss bridge. FBGs FBGs have have also also been been used used for for damage damage detection detection in in large large scale scale structures. structures. In In particular, particular, carbon carbon fiber polymer matrix composites are used in many advanced aerospace applications; this is fiber polymer matrix composites are used in many advanced aerospace applications; this is primarily primarily duehigh to their highstrength specific and strength and in stiffness in the fiber direction. However, even with due to their specific stiffness the fiber direction. However, even with modest modest loading, CFRPs develop complicated internal damage that can include matrix cracking, loading, CFRPs develop complicated internal damage that can include matrix cracking, delaminations, delaminations, and fiber [29]. Pena et al. [30] used surface mounted and FBGs embedded FBGs and fiber breakage [29]. breakage Peña et al. [30] used surface mounted and embedded in a COPV in a COPV and developed a strain-measurement-based SHM parameter that can be used to track and developed a strain-measurement-based SHM parameter that can be used to track the COPV the COPV structural health throughout loading. Also, acoustic emission is often used as a nonstructural health throughout loading. Also, acoustic emission is often used as a non-destructive destructive method for early detectionXiao of damage. et al. [15] demonstrated of FBGs to method for early detection of damage. et al. [15]Xiao demonstrated the use of FBGsthe to use simultaneously simultaneously monitor both strains and acoustic signals for monitoring load and damage in monitor both strains and acoustic signals for monitoring load and damage in aircraft structures. aircraft structures. Mendoza et al. [31,32] reported on the development of a wireless in-flight FBGMendoza et al. [31,32] reported on the development of a wireless in-flight FBG-based acoustic emission based acoustic emission monitoring systemThe known FAESense™. Thesignals strain were variations and AE monitoring system known as FAESense™. strainasvariations and AE used to verify signals were used to verify the occurrence and location of the damage. Betz et al. [33] developed a the occurrence and location of the damage. Betz et al. [33] developed a damage identification system damage identification forFBGs aerospace uses FBGs sense ultrasound by for aerospace structuressystem that uses to sensestructures ultrasoundthat by obtaining thetolinear strain component obtaining the linear strain component produced by Lamb waves. produced by Lamb waves. 4. 4. Determination Determinationofofthe theWing WingShape Shapeand andIn-flight In-FlightLoads Loadsofofan anUltralight UltralightComposite CompositeWing Wing The carbon-compositemulti-tapered multi-tapered wing a sailplane-inspired 2) was The carbon-composite wing of aof sailplane-inspired aircraftaircraft (Figure(Figure 2) was statically statically strain measurements fromused FBGsto were used the to out-of-plane determine the out-of-plane tested andtested strain and measurements from FBGs were determine deflections and deflections and loads [7] using the FOSS system [4,16]. The wing was instrumented with two optical loads [7] using the FOSS system [4,16]. The wing was instrumented with two optical fibers along the fibers alongone theon main spar, one theone upper skin and skin. one on the lower skin. These measurements main spar, the upper skinon and on the lower These measurements were compared to were compared to four surface strain gages and The a displacement sensor. The wing four surface mounted strain gagesmounted and a displacement sensor. wing assembly was subjected to assembly was subjected to distributed loads to simulate an in-flight maneuver using a three-tier distributed loads to simulate an in-flight maneuver using a three-tier whiffletree mechanism that was whiffletree mechanism in aFBG previous study [34]. The resulting FBGtostrain developed in a previous that studywas [34].developed The resulting strain measurements were compared those measurements were compared to those obtained from conventional strain gages. obtained from conventional strain gages.
Figure Figure 2. 2. CAD CAD model model of of aa sailplane-inspired sailplane-inspired air air vehicle. vehicle.
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2016, 3, 18 5 of 15 4.1.Aerospace Description of Composite Wing 4.1. Description of Composite Wing The composite wing has a semi-span of 5.5 m, a root chord dimension of 0.74 m, and a 4.1. Description of Composite Wing The composite wing has of a semi-span 5.5 root m, a root chordFigure dimension of 0.74the m, wing’s and a maximum maximum airfoil thickness 0.1 m atofthe section. 3 shows primary The composite wing has a semi-span of 5.5 m, a root chord dimension of 0.74 m, and a airfoil thickness of 0.1 m at the root section. Figure 3 shows the wing’s primary structures, structures, which include the upper lower a root rib, a fore aft wing’s spar, and awhich main maximum airfoil thickness of 0.1 and m at the skins, root section. Figure 3 spar, showsanthe primary include the upper and lower skins, a rootfrom rib, athe fore spar, an aft spar, and a main (center) spar. The main (center) spar. The main spar protrudes root rib to form the carry-through structure into structures, which include the upper and lower skins, a root rib, a fore spar, an aft spar, and a mainthe spar protrudes fromtest thearticle root rib to form the carry-through structure into the fuselage. The wing test fuselage. The wing does not have aileron or spoiler cut-outs. (center) spar. The main spar protrudes from the root rib to form the carry-through structure into the article not wing have aileron or spoiler Alldoes structural components of the wing fabricated from woven and unidirectional carbonfuselage. The test article does notcut-outs. have are aileron or spoiler cut-outs. AllAll structural components of the wing are fabricated from woven and unidirectional carbon-fiber/ fiber/epoxy prepregs (Toray composites). The upper and lower skins areand sandwich composites with structural components of the wing are fabricated from woven unidirectional carbon® epoxy prepregs (Toray composites). The upper and lower skins are sandwich composites with a low a low density prepregs foam core (DIAB DivinycellThe HT50, Diab AB, Helsingborg, Sweden) of fiber/epoxy (Toray composites). upper andInternational lower skins are sandwich composites with ® HT50, Diab International AB, Helsingborg, Sweden) of thickness density foam core (DIAB Divinycell ® HT50, thickness 3.175-mm; engineering properties of the composite materials are given inSweden) Table 1. of The a low density foamthe core (DIAB Divinycell Diab International AB, Helsingborg, 3.175-mm; the engineering properties of the composite materials areleading givenare inedge, Table 1. foam core thickness 3.175-mm; properties of 40-mm the composite materials giventrailing in The Table 1. The foam core extends overthe theengineering wing surface to about from the edge, and extends over the wing surface to about 40-mm from the leading edge, trailing edge, and side boundaries. foam core extends over the wing surface towere aboutfabricated 40-mm from the leading trailingbefore edge, and side boundaries. All structural components separately and edge, oven-cured being Allside structural components fabricated separately and oven-cured before being bonded boundaries. All components were separately and oven-cured adhesively bonded in structural an were assembly jig. Steel lift fabricated pins were mounted to the rootadhesively rib before prior being to final adhesively bonded in lift an load assembly Steel the lift pins were mounted toas the rootinrib final in an assembly jig. Steel pins were jig. mounted to the root prior to final assembly toprior provide shear assembly to provide shear transfer from wing torib the fuselage, seen Figure 4.toFurther assembly tofrom provide shear load from wing to theetfuselage, seen inregarding Figure 4. Further load transfer the wing to thetransfer fuselage, asthe seen in Figure 4.al.Further the wing details regarding the wing structure can be found in Sullivan [34]. asdetails details regarding the wing structure can be found in Sullivan et al. [34]. structure can be found in Sullivan et al. [34].
Figure components [34]. Figure 3.3.Composite Composite wing structural components[34]. [34]. Figure3. Compositewing wing structural structural components Table 1. Physical and wingmaterials materials[34]. [34]. Table Physical andengineering engineeringproperties properties of of carbon-composite carbon-composite wing Table 1. 1. Physical and engineering properties
Material WovenFabric Fabric Unidirectional Fabric Material Woven Unidirectional Fabric Material Woven Fabric Unidirectional Properties Toray-T700G Toray-T700S Properties Toray-T700G Toray-T700S Properties Toray-T700G Fabric Toray-T700S 11 , GPa 5.54× ×1010 1.19 × 1022 2 E11E, 11 GPa 5.54 1.19 1 E11 , GPa 5.54 ˆ 10 1.19 ˆ 10 1 1 , GPa 5.54 ×1010 9.31 ×× 10 1000 0 E22E, 22 GPa 9.31 1 E22 , GPa 5.54 ×5.54 ˆ 10 9.31 ˆ 10 0 12, GPa 4.21×4.21 ×1010 4.21 × 1000 0 0 G12,GGPa 4.21 G12, GPa 4.21 ˆ 100 4.21׈1010 −2 −1 3.00×3.00 ×1010 3.10 10 −2 ν12 ˆ 10´2 3.10×׈10 10−1´1 ν12ν12 3.00 3.10 3 3 3 3 3 3 ρ, kg/m 1.49×1.49 ×1010 1.52 ׈10 1010 1.52× 3 ρ, kg/m 3ˆ 10 3 ρ, kg/m 1.49 1.52
Foam FoamCore Core(DIAB (DIAB
Foam Core (DIAB ®® Divinycell HT50) Divinycell HT50) ® HT50) Divinycell −2 −2 8.50 8.50× ×1010
8.50 ˆ 10´2 –– – – – – ´1× 10−1 −1 3.20 ˆ3.20 10 3.20 × 10 −1 ´1 4.95 ˆ4.95 10 4.95× ×1010−1
Figure 4. Top view of the spar/spar and wing/fuselage attachment region [34].
Figure 4. 4. Top Top view view of of the the spar/spar spar/spar and Figure and wing/fuselage wing/fuselageattachment attachmentregion region[34]. [34].
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4.2. Experimental 4.2. ExperimentalSetup Setupand andProcedure Procedure The use the the high high density densityofofFBG FBGstrains strainstotoobtain obtainthe the Theprimary primaryobjective objective of of this this study study was to use deflected wing shape and out-of-plane loads forfor a carbon composite wing subjected to to distributed and deflected wing shape and out-of-plane loads a carbon composite wing subjected distributed concentrated loads. loads. A universal test frame shownshown in Figure 5, was used asused the support structure and concentrated A universal test (UTF), frame (UTF), in Figure 5, was as the support allcases. static test cases. forstructure all staticfor test
(a)
(b)
Figure 5. (a) Sketch of universal test frame (UTF) for concentrated and distributed loading with the Figure 5. (a) Sketch of universal test frame (UTF) for concentrated and distributed loading with the wing wing loading stations LS1, LS2, LS3 and LS4 [34] and (b) photo of wing structure under whiffletree loading stations LS1, LS2, LS3 and LS4 [34] and (b) photo of wing structure under whiffletree loading. loading.
4.2.1. 4.2.1.Wing WingLoading LoadingMethodology Methodology AAthree-tier an in-flight in-flightlift liftdistribution distributionononthe thewing wing three-tierwhiffletree whiffletreesystem system was was used used to to simulate simulate an structure. As illustrated in Figure 5, a saddle fixture that simulated the fuselage-wing connection structure. As illustrated in Figure 5, a saddle fixture that simulated the fuselage-wing connection region was wing structure structurefor fortesting. testing.ToTotransfer transferthe the load region wasused usedwith withthe theUTF UTFto tomount mount the the inverted inverted wing load totoeach wereplaced placedatatthe the span-wise loading stations LS2,and LS3,LS4. andThe LS4. eachwing, wing,four four cradles cradles were span-wise loading stations LS1, LS1, LS2, LS3, The and bottom cradles were designed closely match airfoil shape wing. Wood toptop and bottom of of thethe cradles were designed to to closely match thethe airfoil shape of of thethe wing. Wood spacers ofof 25.4-mm thickness were placed under thethe upper cradles at the spar locations to avoid local spacers 25.4-mm thickness were placed under upper cradles at the spar locations to avoid damage to the wing composite wing was sandwiched between the topthe and cradles local damage to theskins. wing The skins. The composite wing was sandwiched between topbottom and bottom prevent slippage of theand cradles and to minimize of theduring wing during loading. The tocradles preventtoslippage of the cradles to minimize twistingtwisting of the wing loading. The cradles cradles were bolted together and attached to the first tier of the whiffletree using steel connectors. were bolted together and attached to the first tier of the whiffletree using steel connectors. The second The second and third levels of consisted the whiffletree consisted of aluminum C-channel sections and an and third levels of the whiffletree of aluminum C-channel sections and an aluminum I-beam, aluminum I-beam, respectively. The load actuator wassurface attached to ausing floor surface plate using a ball respectively. The load actuator was attached to a floor plate a ball bearing to allow the bearing to allow the actuator freedom of rotation during the loading process. The loading was actuator freedom of rotation during the loading process. The loading was applied by manual pumps applied manual pumps were connected to each actuator. that were by connected to each that actuator. The loading sequence, consisting thirteen loading increments included weight of wing, the The loading sequence, consisting of of thirteen loading increments included the the weight of the wing, each cradle, each tier of the whiffletree, and the five load increments of approximately 222-N, each cradle, each tier of the whiffletree, and the five load increments of approximately 222-N, for a for a total load on the wing of 2269-N, is given in Table 2. After each loading increment, five sets of total load on the wing of 2269-N, is given in Table 2. After each loading increment, five sets of strain strain data from the FBG sensors and the strain gages were recorded and the averaged value was data from the FBG sensors and the strain gages were recorded and the averaged value was used in the used in the calculation of the displacement and wing shape at each load. calculation of the displacement and wing shape at each load. Concentrated loads were applied to the wing at a distance of 4.3-m from the root rib, as shown Concentrated loads were applied to the wing at a distance of 4.3-m from the root rib, as shown in in Figure 6. These point loads were applied by placing lead weights into a container that was Figure 6. These point loads were applied by placing lead weights into a container that was suspended suspended from the center of gravity of the wing by a wooden cradle. Data was acquired after each from the center of gravity of the wing by a wooden cradle. Data was acquired after each load increment load increment listed in Table 3. listed in Table 3.
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Table 2. Wing distribution loading profile.
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Load Step Load Increment (N) Total Applied Load (N) Table 2. Wing distribution loading profile. 1 Wing (155.68) 155.68 2 Cradle (142.74) 298.42 Load Step Load LS1 Increment (N) Total Applied Load (N) 3 Cradle LS2 (147.27) 445.69 1 Wing (155.68) 155.68 4 Cradle LS3 (151.68) 597.37 2 Cradle LS1 (142.74) 298.42 5 Cradle LS4 (156.12) 753.49 3 Cradle LS2 (147.27) 445.69 4 CradleTier LS3 1(151.68) 597.37 6 Whiffletree (112.09) 865.58 5 CradleTier LS4 2(156.12) 753.49 1001.25 7 Whiffletree (135.67) 6 Whiffletree Tier 1 (112.09) 865.58 8 Whiffletree Tier 3 (148.94) 1150.19 7 Whiffletree Tier 2 (135.67) 1001.25 9 222.41 8 Whiffletree Tier 3 (148.94) 1150.19 1372.60 10 444.82 9 222.41 1372.60 1591.54 10 444.82 1591.54 1837.85 11 667.23 11 667.23 1837.85 2036.66 12 889.64 12 889.64 2036.66 13 1112.05 13 1112.05 2269.45 2269.45 Table 3. Concentrated loading profile for wing flexural rigidity calculation. Table 3. Concentrated loading profile for wing flexural rigidity calculation.
Load Step Load Step 1 1 2 2 3 3 4 4 5 5 6 6 7 7 8 8
Load Increment (N) Total Applied Load (N) Load Increment (N)Section Total Applied Load183.82 (N) Cradle LS4 + Channel Cradle Container LS4 + Channel Section 183.82 (8.87) 192.69 Container (8.87) 192.69 46.62 239.32 46.62 239.32 46.62 285.94 46.62 285.94 44.40 330.34 44.40 330.34 44.40 374.74 44.40 374.74 44.40 419.14 44.40 419.14 48.84 467.98 48.84 467.98
Figure 6. Concentrated load test of composite wing. Figure 6. Concentrated load test of composite wing.
4.2.2. Wing Instrumentation 4.2.2. Wing Instrumentation Two continuous optical fibers were affixed to the upper and lower surfaces of the left wing Two continuous optical fibers were affixed to the upper and lower surfaces of the left wing using using MasterBond Supreme 33 (Hackensack, NJ, USA) adhesive. The FBG fibers, with a sensor MasterBond Supreme 33 (Hackensack, NJ, USA) adhesive. The FBG fibers, with a sensor spacing of spacing of approximately 12.5-mm, were placed from root to tip along the main spar, resulting in a approximately 12.5-mm, were placed from root to tip along the main spar, resulting in a total of 388 total of 388 sensors on the upper surface and 390 sensors on the lower surface of the composite sensors on the upper surface and 390 sensors on the lower surface of the composite wing. Figure 7a wing. Figure 7a shows the location of the sensors on the wing structure. Four general-purpose shows the location of the sensors on the wing structure. Four general-purpose strain gages (Vishay strain gages (Vishay Measurements CEA-06-125UN-350, Micro-Measurements, Raleigh, NC, USA), Measurements CEA-06-125UN-350, Micro-Measurements, Raleigh, NC, USA), with a nominal gage with a nominal gage length of 3.175-mm, were installed along the main spar, next to the FBG fiber length of 3.175-mm, were installed along the main spar, next to the FBG fiber at the four gage stations at the four gage stations GS0, GS1, GS2, and GS3, shown in Figure 7a. The strain gages were placed GS0, GS1, GS2, and GS3, shown in Figure 7a. The strain gages were placed approximately 12.7-mm approximately 12.7-mm from the optical fiber, as shown in Figure 7b. In addition to the FBG from the optical fiber, as shown in Figure 7b. In addition to the FBG sensors and strain gages on the sensors and strain gages on the surface of the left wing, both the right and left wings were
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surface of the left wing, both the right and left wings were instrumented with a tip displacement gage instrumented with a tip displacement gage (Celesco PT1DC, Chatsworth, CA, USA). As shown in (Celesco PT1DC, Chatsworth, CA, USA). As shown in Figure 5, a 50-kN load cell (Interface Advanced Figure 5, a 50-kN load cell (Interface Advanced Force Measurement Model 1210BXV-50kN, Force Measurement Model 1210BXV-50kN, Scottsdale, AZ, USA) was placed between the last tier of Scottsdale, AZ, USA) was placed between the last tier of the whiffletree mechanism and the thehydraulic whiffletree mechanism and the hydraulic loading cylinder. loading cylinder. ® [35] software was used with a computer data acquisition system to obtain the strains LabVIEW LabVIEW® [35] software was used with a computer data acquisition system to obtain the and applied at each loading increment. Prior to loading wings, the gages were balanced strains andloads applied loads at each loading increment. Prior the to loading thestrain wings, the strain gages to were a zerobalanced mean voltage by adjusting the output voltage from the strain gage amplifiers. Two channels to a zero mean voltage by adjusting the output voltage from the strain gage (one for each surface mounted fiber the surface wing structure) an eight channel FOSS system were used amplifiers. Two channels (one foron each mountedoffiber on the wing structure) of an eight to channel acquire FOSS the data fromwere the 778 sensors. system usedFBG to acquire the data from the 778 FBG sensors.
(a)
(b) Figure 7. (a) Sensor location on the surface of the wing structure and (b) Strain gage location next to Figure 7. (a) Sensor location on the surface of the wing structure and (b) Strain gage location next to optical fiber [7]. optical fiber [7].
4.3. Out-of-Plane Displacements and Loads Algorithms 4.3. Out-of-Plane Displacements and Loads Algorithms The deflection and load algorithms used in this study are based on classical beam theory. It is deflection and load algorithms used inwing this was study are based classical beam theory. It is to to The be noted that the loading of the composite applied in a on manner to minimize all twist, be thereby noted that the loading of the composite wing was applied in a manner to minimize all twist, thereby considering only the case of pure bending. considering only the case of pure bending. 4.3.1. Out-of-Plane Displacements Algorithm 4.3.1. Out-of-Plane Displacements Algorithm Utilizing the surface strain measurements at multiple equally spaced sensing stations the Utilizing the surface measurements at multiple equally spaced sensing stations the displacement equation canstrain be expressed using discrete strain measurements as [36,37]: displacement equation can be expressed using discrete strain measurements as [36,37]: 2
yi “
»
p∆Lq – p3 i ´ 1q ε 0 ` 6 6c
iÿ ´1 j“1
fi pi ´ jq ε j ` ε i fl , pi “ 1, 2, . . . , nq
(3)
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yi =
(ΔL )2 (3 i − 1)ε 6c
0
+6
i −1
(i − j )ε j =1
j
+ εi , (i = 1, 2,, n)
9 of(3) 15
where yyisisthe thedeflection deflection at the strain sensing station i, the ΔL spacing is the spacing strain stations, sensing where at the strain sensing station i, ∆L is betweenbetween strain sensing stations, c is the distance from the neutral axis to the outermost surface, and ε is the strain measured c is the distance from the neutral axis to the outermost surface, and ε is the strain measured at the ith For FBG. For demonstration, 8 shows the elastic ofloaded a tip loaded cantilevered th the iat FBG. demonstration, FigureFigure 8 shows the elastic curve curve of a tip cantilevered beam beam from from classical beam theory and from using Equation (3); using only nine stations (n = 9), excellent classical beam theory and from using Equation (3); using only nine stations (n = 9), excellent agreement agreement is obtained.is obtained.
Figure 8. 8. Displacement curve of of aa cantilevered cantilevered beam beam subjected subjected to to aa tip tip load. load. Figure Displacement curve
4.3.2. Out-of-Plane Out-of-Plane Loads Loads Algorithm Algorithm 4.3.2. The loads loads algorithm, algorithm [16,37], which was to determine the out-of-plane loads the The which was used to used determine the out-of-plane loads from the from in-plane in-plane FBG strains, was developed from the classical beam equation for a prismatic beam. The FBG strains, was originally developed by Richards et al. [38], from the classical beam equation for a bending moment can be written in by terms of each FBG strain i as:plates. In the present prismatic beam and later validated Bakalyar and Jutte [16]sensor on flatlocation monolithic
study, the loads algorithm was used for the first time (EI ) i to ε i determine the out-of-plane loads of an air (4) i = vehicle structure. For completeness, a briefMdevelopment of the algorithm is given. The bending ci moment can be written in terms of each FBG strain sensor location i as: where Mi is the bending moment, εi is the strain, (EI)i is the flexural rigidity, and c is the distance pEIqi εi from the neutral axis to the sensor. The moment station i can also be determined by taking Mi “ at each (4) ci the moment at each longitudinal location (L − i∆L) using:
where Mi is the bending moment, εi is theMstrain, (EI)i is)]the flexural rigidity, and c is the distance (5) i = P [(L - i ΔL from the neutral axis to the sensor. The moment at each station i can also be determined by taking the and P is at theeach applied load. Using Equation (5) in Equation (4), the flexural rigidity (EI)i at each FBG moment longitudinal location (L ´ i∆L) using: station is obtained as: Mi “ PPrpL ∆Lqs (5) [(L ´ )]c i - i iΔL (EI )i = (6) εi and P is the applied load. Using Equation (5) in Equation (4), the flexural rigidity (EI)i at each FBG station is obtained as: where L is the length of the beam, i is the sensor station number , ∆L is the P rpL ´ i ∆Lqs ci (i =1, 2, ···, n) pEIq “ i spacing between strain sensors, c is the distance fromε the neutral axis, and ε is the strain at the(6) ith i station. Equation (6) is used to determine the flexural rigidity of a cantilevered beam by obtaining where L is the length of the beam, i is the sensor station number (i = 1, 2, ¨ ¨ ¨ , n), ∆L is the spacing the strain response for a concentrated load at a known location. The flexural rigidity can be used in between strain sensors, c is the distance from the neutral axis, and ε is the strain at the ith station. Equation (4) to determine the bending moment, which is subsequently used to calculate the shear Equation (6) is used to determine the flexural rigidity of a cantilevered beam by obtaining the strain loads Vi from: response for a concentrated load at a known location. The flexural rigidity can be used in Equation (4) ΔMi to determine the bending moment, which is subsequently used to calculate the shear loads Vi from: Vi ≈ (7) Δ xi ∆ Mi Vi « (7) xi where xi is the FBG strain sensor locations. The ∆ out-of-plane loads Pi are determined using the equilibrium of forces, as: where xi is the FBG strain sensor locations. The out-of-plane loads Pi are determined using the equilibrium of forces, as: Pi “ ´∆Vi . (8)
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4.4. Results and Discussion Strain measurements from 778 FBG sensors were used to compute the wing shape and the out-of-plane load per sensor for the concentrated and distributed loading cases. Concentrated loading was used primarily to enable the calculation of the flexural rigidity of the complex wing at each sensor location for subsequent use in the determination of the out-of-plane loads for all test cases. The composite wing was subjected to a distributed loading to simulate an in-flight loading condition. 4.4.1. Concentrated Loading At each of the 778 FBG sensor locations, the flexural rigidity of the wing section was determined (Equation (6)) and the values computed from each load increment were averaged to obtain the effective flexural rigidity. The strain from the largest concentrated load (468-N) and the effective flexural rigidity (EI)i were used in Equation (4) to determine the bending moment Mi . A linear regression fit of the bending moment data was performed and the results were used to calculate the shear force distribution and subsequently the out-of-plane loads using Equation (8). Table 4 shows the results of the load calculation for an applied load of 468-N; the total computed load is within 1.62% of the applied load. The computed out-of-plane loads for each of the load increments, with one exception, are within 2% of the applied load, as shown in Table 5. Table 4. Wing out-of-plane load calculation for the 467.98-N concentrated load. Location (m)
Strain (10´6 )
c (mm)
EI (N¨ m2 )
Moment (N¨ m)
Shear (N)
Load (N)
Applied Load (N)
Difference (%)
0.13 1.03 1.93 2.82 3.72 4.62
´498.26 ´439.48 ´543.22 ´506.35 ´361.35 ´2.07
49.73 48.44 46.51 43.47 38.38 29.80
245,325.24 167,054.17 106,469.51 63,571.24 38,359.38 30,833.92
2158 1731 1303 875.4 447.8 20.18
475.57 475.57 475.57 457.57 475.57 475.57
0.00 0.00 0.00 0.00 0.00 ´475.57
0.00 0.00 0.00 0.00 0.00 ´467.98
0.00 0.00 0.00 0.00 0.00 1.62
Table 5. Concentrated loads: Comparison of applied load and FGB-based calculated out-of-plane load. Applied Load (N)
Calculated Load (N)
Difference (%)
´192.70 ´239.32 ´285.94 ´330.34 ´374.74 ´419.14 ´467.98
´195.00 ´242.11 ´290.82 ´336.26 ´381.17 ´428.55 ´475.57
1.19 1.18 1.71 1.79 1.72 2.25 1.62
4.4.2. Strains under Distributed Loading Figure 9 shows the measured strain distribution from the FBG sensors and the strain gages at the maximum resultant load of 2269-N. The measured strains from the strain gages are slightly offset from those obtained from the FBGs; this was expected since the strain gages could not be placed directly on the FBGs and were actually mounted approximately 13-mm from the optical fiber (Figure 7b). Figure 9 shows a generally linear strain distribution along the wing span. Elevated strains are seen close to the wing root, where the load rapidly transitions to the main spar. The high spatial density of the strains from the FBGs also revealed some physical details not normally obtained from conventional strain gages. For example, the locations designated by the open circles in Figure 9 were determined to be locations directly under the cradles.
The measured strains strain were gages. used to thelocations elastic designated deflection by curve. Thecircles wing sha obtained from FBG conventional Fordetermine example, the the open a resultant load of 2269-N (2.27-g) shown in Figure 10a; for demonstration, in Figurewhiffletree 9 were determined to be locations directlyis under the cradles. mputed elastic curve is superimposed on the deflected wing in Figure 10b. At this loading, Aerospace 2016, 3, 18 11 of 15 4.4.3.tip Deflections under Distributed Loading wing deflection was measured to be 0.295-m, resulting in a 3.75% difference from The measured FBG strains were 6used to determine the elastic deflection curve. The wing shape mputed deflection of 0.284-m. Table shows the calculated deflections for several load cases, a 4.4.3. Deflections under Distributed Loading at a resultant whiffletree load of 2269-N (2.27-g) is shown in Figure 10a; for demonstration, the predictions are The within 4.2% of the measured data. measured FBG strains were used to determine the elastic deflection curve. The wing shape at computedaelastic curve is superimposed on the deflected wing in Figure 10b. At this loading, the resultant whiffletree load of 2269-N (2.27-g) is shown in Figure 10a; for demonstration, the computed left wing elastic tip deflection was measured to be wing 0.295-m, resulting in loading, a 3.75% curve is superimposed on the deflected in Figure 10b. At this the difference left wing tip from the wasof measured to beTable 0.295-m, resultingthe in a calculated 3.75% difference from the computed deflection computeddeflection deflection 0.284-m. 6 shows deflections for several loadofcases, and 0.284-m. Table 6 shows the calculated deflections for several load cases, and the predictions are within the predictions are within 4.2% of the measured data. 4.2% of the measured data.
Figure 9. Strain distribution on the composite wing for a whiffletree resultant load of 2269-N (2.27g). Figure 9. Strain distribution on the composite wing for a whiffletree resultant load of 2269-N (2.27 g).
Figure 9. Strain distribution on the composite wing for a whiffletree resultant load of 2269-N (2.27g).
(a) (a)
(b)
(b) Figure 10. Wing deflection for a resultant whiffletree load of 2269-N showing the (a) calculated elastic curve using FBG strains and (b) displacement results superimposed on the wing under load. Figure 10. Wing deflection for a resultant whiffletree load of 2269-N showing the (a) calculated elastic
Figure 10. Wing deflection for a resultant whiffletree load of 2269-N showing the (a) calculated curve using FBG strains and (b) displacement results superimposed on the wing under load. elastic curve using Table FBG strains and and (b) displacement results superimposed oncase. the wing under load. 6. Measured calculated tip deflection for distributed load Load (N)Table Measured Tip Deflection (m) tip Calculated Tipfor Deflection (m)load Difference (%) 6. Measured and calculated deflection distributed case. 1373 −0.184 −0.178 3.02 −0.209 2.29 Load (N)1592Measured Tip Deflection (m) Calculated−0.205 Tip Deflection (m) Difference (%)
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Table 6. Measured and calculated tip deflection for distributed load case. Load (N) 1373 1592 1837 2036 Aerospace 2016, 3, 18 2269
Measured Tip Deflection (m)
Calculated Tip Deflection (m)
Difference (%)
´0.184 ´0.209 ´0.241 ´0.265 ´0.295
´0.178 ´0.205 ´0.231 ´0.257 ´0.284
3.02 2.29 4.08 3.23 3.75
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4.4.4. Out-of-Plane Loads under Distributed Loading 4.4.4. Out-of-Plane Loads under Distributed Loading The in-plane FBG strain data obtained from the whiffletree (distributed) loading was used to The in-plane FBG strain data obtained from the whiffletree (distributed) loading was used to determine the out-of-plane loads. The effective flexural rigidity, determined from the concentrated determine the out-of-plane loads. The effective flexural rigidity, determined from the concentrated load case, was used to calculate the bending moment acting on the wing. A third-order polynomial load case, was used to calculate the bending moment acting on the wing. A third-order polynomial regression was used to obtain the moment as a function of the span-wise location; this was regression was used to obtain the moment as a function of the span-wise location; this was subsequently subsequently used to calculate the out-of-plane load per sensor, as shown in Figure 11. Table 7 used to calculate the out-of-plane load per sensor, as shown in Figure 11. Table 7 shows that the shows that the computed FBG-based loads are within 4.2% of the measured applied loads. computed FBG-based loads are within 4.2% of the measured applied loads.
Figure11. 11.Out-of-plane Out-of-planeload loadper perFBG FBGsensor sensorfor forresultant resultantload loadof of2269-N. 2269-N. Figure Table 7. Wing distributed loads: Comparison applied loads and calculated out-of-plane loads. Table 7. Wing distributed loads: Comparison of of applied loads and calculated out-of-plane loads.
Applied (N) AppliedLoad Load(N) (N) FBG-Based FBG-BasedLoads Loads (N) 1370 1355 1370 1355 1589 1641 1589 1641 1835 1859 1835 1859 2033 2116 2033 2116 2269 2364 2269 2364
Difference (%) Difference (%) 1.1 1.1 3.2 3.2 1.3 1.3 4.1 4.1 4.2 4.2
Conclusions 5.5.Conclusions This paper an overview of FBG-based sensing for large scale This paperprovides provides an overview of FBG-based sensing for applications, large scale demonstrating applications, the real-world applications and increasing usage of optical fiber measurement methods. These sensors demonstrating the real-world applications and increasing usage of optical fiber measurement are being used every industry, particularly aerospace andparticularly civil structures, measurement and methods. Theseinsensors are being used in for every industry, for for aerospace and civil monitoring of strains, temperatures and pressures and for detecting damage. The many attributes of structures, for measurement and monitoring of strains, temperatures and pressures and for FBG sensors make them ideal for measurement and monitoring applications. detecting damage. The many attributes of FBG sensors make them ideal for measurement and For demonstration monitoring applications.of an operational application, a fiber optic strain sensing system was used to obtain a strain distribution over the semi-span of a full scale composite wingsensing subjected to concentrated For demonstration of an operational application, a fiber optic strain system was used and distributed The high spatial resolution of of the FBG strain measurements enabled the to obtain a strainloads. distribution over the semi-span a full scale composite wing subjected to computation and of the deflected loads. wing shape and spatial out-of-plane loadsofofthe the FBG wing.strain When compared to concentrated distributed The high resolution measurements conventional foil strain gages, thedeflected FOSS system a much higher density 12.5-mm) of enabled the computation of the wingprovided shape and out-of-plane loads of(every the wing. When strain measurements, thereby detailsthe of the structure under load. a much higher density compared to conventional foilrevealing strain gages, FOSS system provided
(every 12.5-mm) of strain measurements, thereby revealing details of the structure under load. Acknowledgments: The support provided for this study by the NASA Armstrong Flight Research Center (Award No. AERO532 11020161), Raspet Flight Research Laboratory, and the NASA/Mississippi Space Grant Consortium (Award No. 12040456 12070825) is gratefully acknowledged. Author Contributions: This research was conducted under the guidance of Rani W. Sullivan and W. Lance Richards. Matthew J. Nicolas conducted the experimental study as a graduate research assistant at Mississippi
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Acknowledgments: The support provided for this study by the NASA Armstrong Flight Research Center (Award No. AERO532 11020161), Raspet Flight Research Laboratory, and the NASA/Mississippi Space Grant Consortium (Award No. 12040456 12070825) is gratefully acknowledged. Author Contributions: This research was conducted under the guidance of Rani W. Sullivan and W. Lance Richards. Matthew J. Nicolas conducted the experimental study as a graduate research assistant at Mississippi State University. All authors contributed to the writing of this manuscript. Conflicts of Interest: The authors declare no conflict of interest.
Abbreviations FBG NASA UAV FOSS CFRP COPV UTF
Fiber Bragg Grating National Aeronautics and Space Administration Unmanned Aerial Vehicle Fiber Optic Strain Sensing Carbon Fiber Reinforced Plastic Composite Overwrapped Pressure Vessel Universal Test Frame
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