SPECIALTY MATERIALS, INC.

Transcription

1 SPECIALTY MATERIALS, INC. Manufacturers of Boron and SCS Silicon Carbide Fibers and Boron Nanopowder EVALUATION OF BORON/EPOXY DOUBLERS FOR REINFORCEMENT OF COMMERCIAL AIRCRAFT METALLIC STRUCTURES Prepared by: Boeing Company March 1996 Appendices available from Specialty Materials upon request. D

2 CAGE CODE: THIS DOCUMENT IS: CONTROLLED BY: Structures Research and Technology, PREPARED UNDER: CONTRACT NO A3397R4 DOCUMENT NO: D MODEL: All TITLE: Evaluation of Bonded Boron/Epoxy Doublers for Reinforcement of Commercial Aircraft Metallic Structures THE INFORMATION CONTAINED HEREIN IS NOT PROPRIETARY PREPARED BY: Paul Rutherford /15/96 Steve Berg /15/96 CHECKED BY: Chris Mazur B-YH12 4/2/96 APPROVED BY: Matt Miller B-YH12 3/27/96 SIGNATURE ORGN DATE D

3 Executive Summary For a number of years, researchers have been developing a technology for repair and reinforcement of airplane structures using bonded boron/epoxy doublers. This technology has been successfully used on military airplanes such as reinforcement of the F-111 wing pivot fitting, repair of stress corrosion on the C-130, and fatigue cracking on the Mirage El and Macchi airplanes. More recently, extensive application has been seen on the USAF C-141 fleet (wing weep-hole riser cracks). The concept involves bonding a boron/epoxy doubler over the metallic structure containing damage. The boron/epoxy doubler has a higher stiffness than aluminum or titanium and thus transfers load through the adhesive to the composite doubler bypassing the damaged metal structure. This reduces the stress level in the metal structure which will prevent or slow continued damage accumulation. The boron/epoxy doubler reinforcement concept has sparked interest by commercial airlines who are always looking for quicker, more efficient repair techniques. Potential commercial airplane applications of this technology include repairs to damaged secondary and primary structure, and as a preventive stress reliever modification to a problem structural item in the fleet. The Boeing Company was contracted by Textron Specialty Materials through Boeing Technology Services, contract A3397R4, to develop a test program which would provide structural data on bonded boron/epoxy repairs to metallic aircraft structures. This proposed test program was developed under a phase I contract. The phase II contract was tasked with accomplishing the test program proposed in the phase I final report. The test plan and tasks performed in phase II were subsequently modified as data was obtained. The report is separated into three sections: boron/epoxy material and doubler installation process specification, nondestructive inspection, and structural analysis and performance tests. The writing of the installation process specification for bonding boron/epoxy to metallic aircraft structures (with emphasis on aluminum) was accomplished. This involved chemical, physical, and mechanical characterization tests to understand the boron tape with the 250 F epoxy plus numerous process sensitivity studies. It was determined that the boron/epoxy material met AMS 3 867/4A specification requirements when tested to the AMS specification procedures. The material has similar physical properties to other 250 F curing composite material systems; namely BMS materials. The doubler installation cure conditions were actually improved as a result of the process sensitivity studies. D

4 Nondestructive inspection techniques were evaluated for the detection of: disbond or delamination of the doubler, and, crack growth under the doubler. Textron took the lead role in this effort with the development of the NDI procedures while Boeing contributed a consulting role with design of the reference standards and evaluating the NDI procedures. The structural analysis and performance tests were separated into material property testing, stress analysis, and field repair testing. The stress analysis was performed to support all the specimen designs and is referenced in the material property testing and the field repair testing sections. The material property tension testing was accomplished on three batches of boron/epoxy with two different specimen configurations: unidirectional and bow-tie (0 ±45 ). This was done to verify differences between the autoclave process and the heat blanket/vacuum bag process and differences between the 350 F and 250 F cure epoxy resins. The bulk of the testing was performed on the field repair specimen which was a boron/epoxy doubler bonded to a damaged 4x16 inch 7075-T6 thin aluminum sheet. Field repair testing was divided into two separate sections: baseline geometry determination tests and parametric tests. The baseline geometry tests examined different configuration variables relating to the design of the boron/epoxy doubler. The intent in this testing is to culminate in selection of one design for further evaluation in the parametric tests. The parametric tests were designed to examine different variables on the selected baseline design. Some of the variables investigated were doubler ply reduction, aluminum thickness changes, crack length, environmental exposure, different doubler anomalies, doubler impact events, and disbonds. From the parametric tests performed the following were found: the -65 F exposure had a significant impact on the fatigue life of the specimen with a doubler; restraining the lateral bending during the -65 F and room temperature fatigue tests significantly increase the fatigue life of the specimen; effects of fasteners under the doubler and in-line with the flaw had a significant impact on the fatigue life where as 0.5 inch simulated disbonds had little effect on the fatigue life of the specimen; when impacted 0.5 inches from the stop drill the doubler seemed to show an increased fatigue life with impact levels less than 300 in-lbs; fatigue stress levels at 0-18 ksi have a greater impact on fatigue life than fatigue stress levels of 3-20 ksi; and doubler applications on thicker metallic substrates have a large impact on fatigue life. D

5 Abstract This document reports the work accomplished on contract A3397R4 which evaluated boron/epoxy doublers for the application to repair or reinforce commercial aircraft. The report is separated into three sections: boron/epoxy material and doubler installation process specification, nondestructive inspection, and structural analysis and performance tests. The writing of an installation process specification for bonding boron/epoxy to metallic aircraft structures (with emphasis on aluminum) was accomplished. Nondestructive inspection techniques were evaluated for the detection of: disbond or delamination of the doubler, and, crack growth under the doubler. The structural analysis and performance tests are separated into material property testing, stress analysis, and field repair testing. The material property tension testing was accomplished on three batches of boron/epoxy with two different specimen configurations: unidirectional and bow-tie (0 ±45 ). This was done to verify differences between the autoclave process and the heat blanket/vacuum bag process and differences between the 350 F and 250 F cure epoxy resins. The bulk of the testing was performed on the field repair specimen which was a boron/epoxy doubler bonded to a damaged 4x16 inch 7075-T6 thin aluminum sheet. Field repair testing was divided into two separate sections: baseline geometry determination tests and parametric tests. The baseline geometry tests examined different configuration variables relating to the design of the boron/epoxy doubler: ply layup, ply orientation, and doubler geometry. It was the intent in this testing to culminate in selection of one design for further evaluation. The parametric tests were designed to examine different variables on the selected baseline design. Some of the variables investigated were doubler ply reduction, aluminum thickness changes, crack length, environmental exposure, different doubler anomalies, doubler impact events, and disbonds. The results of this program are documented. D

6 Table of Contents List of Figures List of Tables Introduction Background Textron Specialty Materials Contract Boron/Epoxy Material and Doubler Installation Process Specification Baseline Material Characterization Chemical Characterization Infrared Spectroscopy (IR) High Pressure Liquid Chromatography (HPLC) Physical Characterization Resin Content, Uncured Prepreg Volatile Content Resin Flow Gel Time Autoclave-cured Laminate Ply Thickness Laminate Fiber Volume/Void Content Flammability Mechanical Characterization Compression Strength Short Beam Shear Strength Tension Strength Single Overlap Shear Strength Wedge Crack Growth Flatwise Tension Strength Conclusions Process Sensitivity Studies Cure Temperature Cure Pressure Heat-up Rate Sensitivity Anodize Parameters Primer Cure Conditions Fluid Exposure of Cured Coupons Conclusions D

7 2.3 Installation Process Specification Specification Development Specification Validation Fabrication of Field Repair Test Specimens Nondestructive Inspection (NDI) Ultrasonic Resonance Eddy Current Structural Analysis and Performance Tests Material Property Tension Tests on Vacuum-Bag Cured Boron/Epoxy Environmental Testing Unidirectional Tensile Specimen Tests Bow-Tie (0 +45 ) Tensile Specimen Tests Stress Analysis INCAP Laminate Analysis COSMOS 2D Finite Element Analysis NIKE 3D Finite Element Analysis ANSYS 2D Finite Element Analysis Field Repair Tests Baseline Geometry Tests Static Tests Fatigue Tests Impact Screening Tests Parametric Tests Additional Tests on the Baseline Geometry Change in Laminate Thickness Crack Size in Different Aluminum Thicknesses Environmental Exposure Effects of Fasteners and Disbonds Impact Stop Drilling Crack Tip or Not Changes in Ply Drop-off Rate Changes in Cure Pressure Restraining Lateral Bending Changes to Fatigue Spectrum Changes in Fatigue Load Level inch Aluminum Investigation D

8 4.3.3 Microscopic and Failure Analysis Summary Appendix: A Process Specification for the Fabrication and Application of Boron/Epoxy Doublers to Aluminum Structure... Al B Environmental Conditioning Information... Bl C Material Property Results... Cl D Field Repair Test Matrix... Dl E INCAP Laminate Analysis... El F COSMOS Finite Element Analysis... Fl G NIKE Finite Element Analysis... Gl H ANSYS Finite Element Analysis... HI I Field Repair Drawings... II J Through-Transmission Ultrasonic (TTU) Nondestructive Inspection (NDI)......Jl K Field Repair Results - Baseline Geometry Testing... Kl L Field Repair Results - Parametric Testing... LI M Microscopy and Failure Analysis... Ml List of Figures Figure Fabrication Procedure for Field Repair Test Specimens...61 Figure Boron/Epoxy Nondestructive Inspection Reference Standard for Disbond and Delamination Evaluation Figure Boron/Epoxy Reference Standard Eddy Current NDI for Crack Growth Detection Below the Doubler Figure Material Property Values for Different Boron/Epoxy Tapes 70 Figure Material Property Tension Test Matrices...71 Figure Material Property Values Published in the NAS A/DOD Composite Design Guide...72 Figure Unidirectional Ply Tensile Test Specimen...74 Figure Photograph of Test Setup for Unidirectional Tension Tests..74 Figure Photograph of Tab Failure on Unidirectional Tension Specimen Figure Bow Tie Tensile Test Specimen...75 Figure Photograph of Test Setup for Bow Tie Tension Tests...76 D

9 Figure Summary of the Material Property Results for the Bow Tie Tension Tests - Tensile Strength 77 Figure Summary of the Material Property Results for the Bow Tie Tension Tests Tensile Modulus 77 Figure Photograph of Typical Failure in Gage Section of a Bow Tie Tension Specimen...77 Figure Shear Stress vs. Shear Strain for American Cyanamid FM73 Film Adhesive...79 Figure BMS Film Adhesive Material Properties Reported from Vendors...79 Figure Generalized Single Shear Lap Joint Stress...80 Figure Normalized Shear Stress vs. Distance Along a Single Shear Lap Joint in Tension Figure Normal and Inverted Ply Lay-ups Figure Lay-up Configurations and Doubler Geometries for Field Repair Static and Fatigue Specimens...84 Figure Numbering System for Fatigue Specimens with Lay-up Configurations and Doubler Geometries Figure Specimen Dimensions for Field Repair Static and Fatigue Tests...86 Figure Location of Instrumentation for Field Repair Static Tests..86 Figure Photograph of Field Repair Static Test Setup...86 Figure Representative Load vs. Axial Strain Plot for a Field Repair Static Specimen Figure Summary of Field Repair Static Results Ultimate Tensile Stress Figure Summary of Field Repair Static Results Global Tensile Modulus Figure Indication if Mechanical Hysteresis has Occurred From Strain Surveys Done During Fatigue Testing...89 Figure Photograph of Setup for Field Repair Fatigue Test with Crack Detection Gage D

10 Figure Photograph of Setup for Field Repair Fatigue Test with Crack Propagation Gages...91 Figure Impact Details of Field Repair Impact Specimens 94 Figure Top View of Three Field Repair Specimens Impacted at 100, 500, and 1200 inch-lbs Figure Bottom View of Three Field Repair Specimens Impacted at 100, 500, and 1200 inch-lbs Figure Side View of Three Field Repair Specimens Impacted at 100, 500, and 1200 inch-lbs Figure Crack Growth Curve for Specimen TM Figure Crack Growth Curves for Specimens TM4.5-1 and TM Figure Crack Growth Curves for Specimens TM5-16 through TM Figure Crack Growth Curves for Specimens TM5-7, TM5-8 and TM5-10 through TM Figure Crack Growth Curves for Specimens TM7-5, TM7-6, TM7-8,TM7-12,TM7-21andTM Figure Crack Growth Curves for Specimens TM7-19 and TM Figure Specimen Configuration B-l Figure Specimen Configuration B Figure Specimen Configuration B Figure Summary of Impact Residual Static Results Ultimate Tensile Stress Figure Summary of Impact Residual Static Results Global Tensile Modulus Figure Crack Growth Curves for Specimens TM13-1, TM13-2, andtm Figure Crack Growth Curves for Specimens TM14-1 and TM Figure Crack Growth Curve for Specimen TM Figure Photograph of Lateral Bending Restraint for Test Matrix 18 Room Temperature Specimens D

11 Figure Photograph of Lateral Bending Restraint for Test Matrix 18, -65 F Specimens Figure Crack Growth Curves for Specimen TM 18-3 and TM Figure Hz Fatigue Test Spectrum Loading Figure Crack Growth Curves for Specimens TM20-2 and TM Figure Crack Growth Curves for Specimens TM21-4 and TM Figure Crack Growth Curves for Specimens TM21-7, TM21-8 andtm List of Tables Table Boron/Epoxy Characterization Tests...17 Table Summary of Program Boron/Epoxy Lots...18 Table Relative HPLC Peak Ratios...19 Table Resin Content, Uncured Prepreg...20 Table Volatile Content of Boron/Epoxy Tape...21 Table Resin Flow Percent of Boron/Epoxy Tape...22 Table Gel Time of Boron/Epoxy Tape...22 Table Cured Laminate Ply Thickness of Boron/Epoxy Tape...23 Table Flammability Properties of Boron/Epoxy Tape...24 Table Baseline Mechanical Characterization Tests...25 Table Baseline 0 Compression Strength, -65 F...26 Table Baseline 0 Compression Strength, Room Temperature...28 Table Baseline 0 Compression Strength, 180 F/Wet...30 Table Baseline Short Beam Shear Strength, Room Temperature.34 Table Baseline Short Beam Shear Strength, 180 F/Wet...35 Table Tension Strength Values, Room Temperature...37 Table Baseline Single Overlap Shear Strength, Room Temperature 38 Table Wedge Crack Growth, PACS Surface Treatment Validation.40 D

12 Table Overlap Shear Strengths for Alternate Cure Time/Temperatures, RT...42 Table Overlap Shear Strengths for Alternate Cure Pressures, RT...45 Table Overlap Shear Strengths for Alternate Heat-up Rates, RT...49 Table Wedge Crack Growth, Elevated Anodize Temperatures...51 Table Overlap Shear Strength, Alternate Primer Cures, RT...53 Table Compression Strength, Fluid Exposure, RT...56 Table Boron/Epoxy Reference Standards Required for Disbond of the Adhesive and Delamination of the Composite...66 Table Tested Impact Energy for Specimens in Figures , , and Table Results of Crack Initiation Sites Analyzed on Field Repair Specimens D

13 1.0 Introduction 1.1 Background For a number of years, researchers have been developing a technology for repair and reinforcement of airplane structures using bonded boron/epoxy doublers. This technology has been successfully used on military airplanes such as reinforcement of the F-l 11 wing pivot fitting, repair of stress corrosion on the C-130, and fatigue cracking on the Mirage HI and Macchi airplanes. Many of these doublers have been in service for 20 years with no reported problems. More recently, extensive application has been seen on the USAF C- 141 fleet (wing weep-hole riser cracks). The concept involves bonding a boron/epoxy doubler over the metallic structure containing damage. The boron/epoxy doubler has a higher stiffness than aluminum or titanium and thus transfers load through the adhesive to the composite doubler bypassing the damaged metal structure. This reduces the stress level in the metal structure which will prevent or slow continued damage accumulation. The boron/epoxy doubler reinforcement concept has sparked interest by commercial airlines who are always looking for a quicker, more efficient repair techniques. Potential commercial airplane applications of this technology include repairs to damaged secondary and primary structure, and as a preventive stress reliever modification to a problem structural item in the fleet. To date, the use of this technology for commercial airplane applications has been limited. Demonstration doublers have been installed over a section of fuselage lap joint on an Australian Airline 727 and at nine locations on a Qantas 747. Further, a boron/epoxy repair was made to a corroded keel beam on an Ansett 767. All of the aforementioned repairs were designed and installed by Australian Aeronautical Research Laboratory (ARL) personnel. A part of this contract, which will not be discussed in this report, was the installation of boron/epoxy doublers over undamaged structural areas on two Federal Express 747's. This was performed to test the durability of the installation process specification developed on this contract. 1.2 Textron Specialty Materials Contract The Boeing Company was contracted by Textron Specialty Materials (hereafter referred to as Textron) through Boeing Technology Services (BTS) to develop a test program which would provide structural data on bonded boron/epoxy repairs to metallic aircraft structures. This proposed test program was developed under a phase I contract. The phase II contract was tasked with accomplishing the test program proposed in the D

14 phase I final report. The test plan and tasks performed in phase II were subsequently modified as data was obtained. The sections of this document report the tasks accomplished on the phase II contract. There were four basic tasks: (1) material properties tests on three batches of boron/epoxy to verify differences between the autoclave process and the heat blanket/vacuum bag process; (2) writing of an installation process specification for bonding boron/epoxy to metallic aircraft structures (with emphasis on aluminum); (3) stress analysis to, design the specimens tested, understand the load transfer mechanisms from the aluminum through the adhesive to the doubler, and approximate the residual thermal stress after cure due to different coefficients of thermal expansion; and, fatigue and static tests on boron/epoxy doublers bonded to 7075-T6 aluminum alloy sheets which had a simulated crack (these tests are called "field repair" tests in this report). 2.0 Boron/Epoxy Material and Doubler Installation Process Specification The objective of this effort was to verify and characterize the material properties of 3 batches of boron/epoxy unidirectional tape and compare them against selected requirements of the AMS 3867/4A and BMS material specifications, then develop an application process specification for the material. The process specification was required to control the application of the 250 F-curing boron/epoxy material onto aluminum aerospace structures. The process specification D titled "Process Specification for the Fabrication and Application of Boron/Epoxy Doublers to Aluminum Structure" is a deliverable for this program. The specification provides the detailed requirements and procedures to install high quality bonded boron/epoxy doublers with a primary focus on aircraft and in field-level maintenance environments. The specification was completed after performing a thorough characterization of the boron/epoxy unidirectional tape and structural bonding materials, extensive investigation of common variables encountered in field-level aircraft maintenance operations, and validation of the specification controls on test specimens. The studies leading to the completion of the specification are detailed in this section. The resultant process application specification is Appendix A. 2.1 Baseline Material Characterization D

15 The material characterization was conducted to understand how the boron/epoxy tape typically performs when processed using the traditional clean room handling and layup procedures followed by autoclave curing. The characterization was also used to establish that all tape was representative of material procured to the AMS 3867/4A specification. The boron/epoxy was subsequently evaluated for performance when processed using typical on-aircraft maintenance equipment (portable processing equipment that relies primarily on vacuum pressure and heat blankets for cure). In addition, the film adhesive used for the doubler bond was also evaluated for mechanical performance and compatibility with the boron/epoxy material. The boron/epoxy was chemically, physically, and mechanically characterized using a number of techniques. The extensive characterization of the boron/epoxy was required because of Boeing's limited working knowledge of the material. Evaluation of the boron/epoxy tape was based on techniques used for 250 F-curing graphite/epoxy materials in accordance with BMS (Boeing Material Specification) "Epoxy Impregnated Graphite Tapes and Woven Fabrics F (121 C) Cure." BMS specifications contain requirements and evaluation procedures similar to many other industry composite material specifications. Some differences exist between the test procedures of the BMS and AMS specifications though no attempt was made during this program to compare the test procedures and identify specific differences. The adhesive was only characterized for mechanical performance using portable processing equipment since the material has been thoroughly evaluated in the course of qualification and use in The Boeing Company under the controls of BMS and the appropriate autoclave processing specifications. The epoxy resin impregnated boron filament tape is controlled by AMS (Aerospace Material Specification) 3 867/4A. At present Textron is the only specification-qualified supplier of this material. The AMS specification contains many material tests and associated requirements that were included in the baseline boron/epoxy characterization, though other characterization tests were conducted in addition to the AMS requirements. The AMS specification tests are listed in Table below along with the program characterization tests. D

16 Table Boron/Epoxy Characterization Tests AMS Spec Tests Program Tests Chemical IR Spectroscopy HPLC Chromatography Physical Volatile Content Volatile Content Resin Solids Resin Flow Gel Time Resin Content Resin Flow Gel Time Tack Laminate Fiber Volume Cured Ply Thickness Laminate Fiber Volume Cured Ply Thickness Laminate Flammability Glass Transition Temp. D

17 Table Boron/Epoxy Characterization Tests (cont.) Mechanical AMS Spec Tests 0 Compression, RT 0 Compression, 180 F 0 Flexure, RT 0 Flexure, 180 F 90 Flexure, RT 90 Flexure, 180 F Program Tests 0 Compression, RT 0 Tension, RT 0 Tension, RT 0 Tension, 180 F 90 Tension, RT 90 Tension, 180 F Short Beam Shear, RT Short Beam Shear, 180 Short Beam Shear, RT 0 Compression, -65 F 0 Comp., 180 FWet SBS, 180 FWet The AMS specification also contains some requirements concerning the boron filaments and the workmanship of the prepreg tape not listed here. The 5521F/4 prepreg tape delivered with specification compliance certifications by Textron for use on this program was not inspected for compliance to the AMS specification requirements by Boeing Receiving Inspection (incoming QA). D

18 Textron delivered 2 shipments of material to Boeing for use on the program. The first shipment was in accordance with the contract proposal and contained 3 batches of material totaling 1280 linear feet of 6 inch wide unidirectional tape as indicated in Table After the initial shipment of material was used up, a second shipment of 900 linear feet was obtained to fabricate the remaining test specimens. Table Summary of Program Boron/Epoxy Lots Batch No. Rolls Linear Feet A A A A All material was maintained in refrigerated storage below 0 F until drawn for use in the test specimens. Out-time for each roll was carefully tracked to assure that the 7 days at 77±5 F specified in the AMS 3867/4A material specification was not exceeded Chemical Characterization Chemical characterization was performed on tape batches A128, A129, and A130 to determine the relative uniformity of the epoxy impregnating resin. Two standard instrumental chemistry techniques, infrared spectroscopy (IR) and high performance liquid chromatography (HPLC), were used in the evaluation Infrared Spectroscopy (JR.) Infrared spectroscopy was performed in accordance with the requirements of BMS The epoxy impregnating resin was extracted from the prepreg using acetone, the resin was transferred to a salt block, and the sample was evaluated in the spectrometer. Two iterations were made for each material batch. The IR data indicates that the resin chemistry is very uniform between the three batches. Data plots are not included with this report because the plotting ink is too faint to reproduce effectively. D

19 High Performance Liquid Chromatography (HPLC) High performance liquid chromatography was performed in accordance with the requirements of BMS which specifies testing per BSS (Boeing Specification Support) 7305 "High Performance Liquid Chromatography - Reverse Phase Method." This specification is similar to others used in the industry but is written specifically for Boeing equipment. Two iterations were made for each material batch. The HPLC data showed good uniformity among the resin in the tape batches. The largest peak in the plots has a retention time of approximately 14.8 minutes. The HPLC data has been ratioed to the largest peak and presented in Table Table Relative HPLC Peak Ratios. Ratio of Peak Area to Peak at 14.8 Minutes Retention Time Retain Time, (min.) A128 Run 1 A128 Run 2 A129 Run 1 A129 Run 2 A130 Run 1 A130 Run The typical method of using HPLC data is to ratio the peak area for curative to the peak area for primary epoxy and similarly for reaction product to primary epoxy which indicate relative amounts of reaction in the staged epoxy impregnating resin. The representative peak for primary epoxy, hardener, and reaction product was not known for this exercise, therefore the ratios presented in Table may not be usable in the traditional D

20 sense. However, more importantly, the consistency of the ratios among the 3 tape batches do indicate the desired chemical uniformity in the impregnating resin. The data of batch ABO run 1 has peaks at all the listed retention times whereas all other batches and runs had one or more of these peaks missing. The presence or absence of these trace peaks is not significant in this evaluation Physical Characterization Tape batches A128, A129, and A130 were characterized for typical physical properties of the boron/epoxy prepreg including resin content of the uncured tape, volatile content of the resin, resin flow, resin gel time, cured laminate ply thickness, and cured laminate fiber volume/void content. A sample from tape batch A130 was further evaluated for flammability characteristics. The boron/epoxy tape compares quite similarly to BMS in the properties characterized. Areas of each tape were also visually inspected for appearance and gaps between the filaments in the tape. All samples had good appearance and no large gaps between the tape filaments: Where possible, values from published literature and similar material specifications are compared to results of the boron/epoxy tape. The physical characterizations were performed using existing Boeing methods and did not use test methods, if they differed, shown in the AMS 3867/4A material specification. These characterizations do not constitute inspection of the measured attributes but rather provide a basic understanding of the boron/epoxy tape Resin Content. Uncured Prepreg Prepreg samples were tested for resin content of the uncured boron/epoxy tape in accordance with the requirements of BMS The test values are shown in Table Table Resin Content, Uncured Prepreg Boron/Epoxy Tape Batch A128 A129 A130 Run 1, % Run 2, % D

21 Run 3, % Average, % Each batch of tape fell within the values published in the Textron literature (33±2%) and also met the requirements of the AMS specification for resin solids (32+2%) Volatile Content Prepreg samples were tested for volatile content in accordance with the requirements of BMS The test values are shown in Table Table Volatile Content of Boron/Epoxy Tape Boron/Epoxy Tape Batch A128 A129 A130 Run 1, % Run 2, % Run 3, % Average, % The boron/epoxy tape samples had extremely consistent values for volatile content. The test values met the Textron published value (1.5% max.) but the values did not meet the requirement of the AMS specification ( %). The low values with regard to the AMS requirements may be due to test differences between the BMS and AMS specifications. No attempt was made to compare the test methods because the results were not critical to our evaluations. The prepreg tape also met the requirements of the BMS specification (2% max.) that govern composite material systems with similar impregnating epoxy resins Resin Flow Prepreg samples were tested for resin flow in accordance with the requirements of BMS The test values are shown in Table D

22 Table Resin Flow Percent of Boron/Epoxy Tape Boron/Epoxy Tape Batch A128 A129 A130 Run 1, % Run 2, % Run 3, % Average, % Two of the three batches of boron/epoxy tape samples met the Textron published value (14+5%) and the requirement of the AMS specification (9-18%). Tape batch A128 did not meet the minimum values of the literature or specification, however the test value could be due to differences between the AMS and BMS specifications. The results were not critical to our evaluations. All batches of the prepreg tape are within the requirements of the BMS specification (7-19%) that govern composite tape material systems with similar impregnating epoxy resins Gel Time Prepreg samples were tested for gel time in accordance with the requirements of BMS The test values are shown in Table Table Gel Time of Boron/Epoxy Tape Boron/Epoxy Tape Batch A128 A129 A130 Run 1, min Run 2, min Run 3, min Avg., min D

23 Neither the Textron literature nor the AMS specification list distinct values for gel time to be used for comparison. All batches of the prepreg tape are within the requirements of the BMS specification (3-13 minutes) that govern composite tape material systems with similar impregnating epoxy resins Autoclave-cured Laminate Ply Thickness Prepreg samples were tested for cured laminate ply thickness in accordance with the requirements of BMS ply test panels were fabricated from the boron/epoxy tape and autoclave cured using the following cure cycle: 5 F/minute heat-up rate to 250±10 F, 85 psig pressure applied and the bag vented at 140 F, a 250 F hold for 90±15 minutes, then ramp down to room temperature. These laminates were also used in later mechanical test coupons. The test values are shown in Table Table Cured Laminate Ply Thickness of Boron/Epoxy Tape Boron/Epoxy Tape Batch A128 A129 A130 Thickness #1, in Thickness #2, in Thickness #3, in Thickness #4, in Thickness #5, in Thickness #6, in Thickness #7, in Thickness #8, in Thickness #9, in Thickness #10, in Avg. panel thick., in Ply thickness, in The cured laminate ply thickness of batch A128 was within the AMS specification requirements of.0052±.0003 inches. Tape batches A129 and A130 were above the D

24 maximum thickness requirement. The cured ply thickness is sensitive to the bagging technique used for cure. No attempt was made to change the bagging technique or fabricate new panels for test since this information was not critical to doubler applications, which incidentally have average ply thicknesses of inch using vacuum cures Laminate Fiber Volume/Void Content Samples from batches A128, A129, and A130 cured using the autoclave cure cycle were tested for laminate fiber volume and void content in accordance with the requirements of BMS The BMS technique calls for a nitric acid digestion cycle to remove the epoxy resin. Unfortunately, the boron fibers also were digested by the nitric acid thus yielding no data. The process was repeated using the optional sulfuric acid for digestion. The sulfuric acid also broke down the boron fibers and no data was obtained. The fiber break down is probably due to differences in test procedure used by Boeing where the sample is heated to much greater temperatures than the 400 F used by Textron in their procedure. The fiber volume/void content data were not pursued further using other techniques. While the data is a good benchmark of laminate quality, fiber volume/void content determinations were not critical to an effective characterization of the material system Flammabilitv A 12-ply 0 unidirectional laminate fabricated from prepreg batch A130 and using the autoclave cure cycle was tested for flammability in accordance with BSS 7230 Method Fl, commonly called the 60 second-vertical test. Results are shown in Table below. Table Flammability Properties of Boron/Epoxy Tape Extinguish Time, (sec.) Burned Length, (in.) Drip Exting. Time, (sec.) Requirement Run No Drip Run No Drip Run No Drip D

25 The boron/epoxy samples easily met the requirements for BMS in the 60 second-vertical test. BMS also has a requirement for the 30 second-45 evaluation which was not performed since the 45 test uses much more material and was not necessary for this characterization Mechanical Characterization The mechanical property characterization is concerned with determining baseline properties for the boron/epoxy as well as for the adhesive used to bond the boron/epoxy to the aluminum substrate and, most importantly, the aluminum substrate surface preparation. The historical difficulties with the application and effectiveness of any bonded doubler to metallic structure has primarily been attributed to the quality of the adhesive bond between the substrate and the doubler and few, if any, problems with the doubler itself. Therefore mechanical characterization tests of Table included bond characteristics to determine baseline material properties. Table Baseline Mechanical Characterization Tests Material/Process Mechanical Test Test Temp., ( F) Boron/Epoxy 0 Compression Strength -65, RT, 180/wet Short Beam Shear Strength RT, 180/wet 0 Tension Strength RT Film Adhesive Single Overlap Shear RT Aluminum Anodize Wedge Crack Growth RT, Salt Spray D

26 The values obtained from the baseline mechanical characterization provide the benchmark against which the subsequent process development and sensitivity evaluations can be effectively compared Compression Strength 0 compression strength properties were determined for 8-ply test coupons in accordance with the requirements of BMS (ASTM D695).\ Coupons were fabricated from material batches A128, A129, and A130. One panel supplied from Textron was also tested at room temperature (RT). Boron/epoxy test panels were fabricated from the tape and autoclave cured using the following cure cycle: 5 F/minute heat-up rate to 250±10 F, 85 psig pressure applied and the bag vented at 140 F, a 250 F hold for 90±15 minutes, then ramp down to room temperature. The coupons were tested at -65 ±5 F dry, room temperature (77±5 F) dry, and 180±5 F after exposure for 2 weeks to 160 F and 95% relative humidity. Coupon strengths are summarized in Tables through for -65 F, RT, and hot/wet tests, respectively. All reported strength values are nominal strengths, that is, based on the nominal thickness values for the coupons. Table Baseline 0 Compression Strength, -65 F Lot A128 Coupon I.D. Width, Thickness, Ult. Ult. Stress (in.) (in.) Load (ksi) (lb.) Stress (ksi) Average= 441 St. Dev.= 16.2 D

27 A Average= 446 St. Dev.= A Average= 444 D

28 St. Dev.= The baseline compression strength properties for boron/epoxy tape at -65 F are similar to the room temperature properties, indicating that the matrix resin is not adversely embrittled by the low temperature. The consistency between the low and room temperature test data is desirable in composite materials. The low individual value in batch A129 would have required a retest if it had occurred in product acceptance testing, however, no retest was performed here as the individual value was not believed to be representative of the material and the purpose of these tests was to understand basic material capabilities which was accomplished with this data. Table Baseline 0 Compression Strength, Room Temperature Lot Coupon Width, Thickness, (in.) Ult.Load Ult. Stress (ksi) Stress (ksi) A128 I.D. (in.) (lb.) Average= 455 D

29 St. Dev.= A Average= 463 St. Dev.= D

30 A Average= St. Dev.= Textron Fabricated Panel Average= 445 St. Dev.= 34.7 The baseline compression strength properties for boron/epoxy tape at room temperature (77±5 F) are consistent with the published literature (425 ksi) and easily meet the 375 ksi average strength requirement of the AMS 3867/4A specification. The Textron panel was tested to confirm that similar values are obtained from the material when processed at 2 different sites using different equipment and technicians. Table Baseline 0 Compression Strength, 180 F/Wet D

31 Lot CouponI.D. Width, (in.) Thickness, (in.) Ult. Load (lb.) Ult. Stress (ksi) Stress (ksi) A Average= 295 St. Dev.= A Average= 247 St. Dev.= D

32 A Average= 239 St. Dev.= The baseline compression strength properties for boron/epoxy tape tested under hot/wet conditions begin to show some scatter among the 3 batches, particularly the 20% higher performance of batch A128. Nevertheless, all batches easily surpassed the AMS 3867/4A average strength requirement of 140 ksi for material tested at 180 F D

33 dry even though the material was tested in the wet condition which is considered to be more severe Short Beam Shear Strength Short beam interlaminar shear strength was determined on autoclave-cured 20-ply 0 test specimens in accordance with the requirements of ASTM D2344. Coupons were fabricated from material batches A128, A129, and A130. Boron/epoxy test panels were fabricated from the tape and autoclave cured using the following cure cycle: 5 F/minute heatup rate to 250±10 F, 85 psig pressure applied and the bag vented at 140 F, a 250 F hold for minutes, then ramp down to room temperature. Coupons were tested at room temperature (77±2 F) and 180 F after 2 weeks (336 hours) exposure in 160 F/95% relative humidity environment. All reported strength values shown in Table and are actual strengths, that is, based on the average measured thickness values for each group of coupons. The baseline short beam interlaminar shear strengths for the boron/epoxy tape measured at room temperature were consistent among the batches. The values obtained were consistent with published literature (14.1 ksi) and meet the strength requirements (13.0 ksi) of the AMS 3867/4A specification. D

34 Table Baseline Short Beam Shear Strength, Room Temperature Coupon Width, Avg.Thick, Ult. Load Ult. Stress Stress I.D. (in.) (in.) (lb.) (ksi) (ksi) Average= St. Dev.= Average= St. Dev.= Average= St. Dev.= D

35 Table Baseline Short Beam Shear Strength, 1'80 F/Wet Coupon Width, Avg.Thick, Ult. Load Ult. Stress Stress Lot I.D. (in.) (in.) (lb.) (ksi) (ksi) A Average= St. Dev.= A Average= St. Dev.= A Average= St. Dev.= Table Baseline Short Beam Shear Strength, 1'80 F/Wet The baseline hot/wet short beam interlaminar shear strengths for the boron/epoxy tape measured at 180 F showed good retention of properties. The values obtained exceeded the AMS 3867/4A specification strength requirements (5.0 ksi) for tape tested at 180 F dry, which is a less severe condition than evaluated here. D

36 Tension Strength 0 tension strength properties were determined for 8-ply test coupons in accordance with the requirements of BMS Coupons were fabricated from material batches A128, A129, and A130, and tested at room temperature (RT). The initial set of tensile coupons failed at approximately 120 ksi average strength, which is less than 50% of expected strengths. A post-test inspection of the failed coupons revealed that all coupons had failed at the termination of the grip tabs. Further inspection of the grip tabs showed that the termination of the tab bevel into the gage section of the specimens was not correctly faired into the specimen. New specimens were fabricated with correctly beveled tabs and subsequently tested. Nominal ply thickness commonly used for many specifications is used for one set of calculations; that is the nominal per ply thickness requirement multiplied by the number of plies as opposed to the actual specimen thickness. Coupon strengths are summarized in Table D

37 Table Tension Strength Values, Room Temperature Coupon I.D. Thickness, (in.) Width, (in.) Ultimate Load, Actual Strength, Nominal Thick., Nominal Strength, (ksi) (in.) (ksi) (lb.) Average Standard 10.2 Deviation 10.1 The test values of Table are still not valid. The tensile values obtained from the retest were improved but still only about 70% of expected values. The material acceptance test values-at Textron for these batches were approximately 230 ksi minimum average D

38 tension strength. Visual inspection of the failed coupons indicated that some coupons did fail in the gage section but many coupons had an unusual out of plane bulge near one failed end of the specimen. Past experience with 0 tension coupons fabricated from other composite materials indicated that extreme care and careful processing controls are required to accurately determine 0 tension properties for many materials. These test problems are related to fiber diameters, hardness, and moduli among other factors of which the combination of boron fiber characteristics increases the test difficulty. Textron has a customized test procedure to overcome these difficulties. Further failure analysis or testing was not conducted as the purpose of this test was to develop baseline properties and not develop proper testing procedures. All further unidirectional material property tension testing for this program was performed at Textron Single Overlap Shear Strength Single overlap shear strength for adhesively bonded aluminum test coupons were determined in accordance with the requirements of BSS (Boeing Support Standard) 7202 Type I. Test panels were fabricated using a single ply of BMS Type II Grade 10 (FM-73M) adhesive applied to aluminum test blanks that were phosphoric acid anodized per B AC 5555 and primed per BAC (BR-127 primer). The test panels were bagged per the BAC processing specification then cured with portable processing equipment and using the following cure cycle: 25 inches Hg vacuum pressure applied, 5 F/minute heat-up rate to 250±10 F, a 250 F hold for 90±15 minutes, then ramp down to room temperature. Individual test coupons were then sectioned from the panels and tested at room temperature (77±5 F). The test values are shown in Table The baseline tensile shear strength values for the coupons cured using standard portable processing methods are slightly below the minimum BMS strength requirements for specimens when cured using autoclave processing (4375 psi average, 4200 psi minimum individual). Inspection of the failed coupons indicated that the adhesive did not have excessive voids generated in the bond line due to the vacuum pressure processing; a common problem for composite materials under vacuum pressure. Table Baseline Single Overlap Shear Strength, Room Temperature D

39 Coupon I.D. Overlap Length Overlap Width Ult. Loa d (lb.) Ult. Stress (psi) Avg.Stress (psi) (in) (in) Wedge Crack Growth Wedge crack growth coupons were fabricated in accordance with the requirements of BAC 5514 to validate the quality of the phosphoric acid anodize surface treatment provided by the portable phosphoric acid anodize containment system (PACS). Two 6 x 6 x inch 7075-T6 aluminum panels were phosphoric acid anodized at room temperature (77±5 F) per the D process specification then primed per BAC Test panels were fabricated using a single ply of BMS Type II Grade 10 (FM-73M) adhesive applied to aluminum test blanks that were phosphoric acid anodized per BAC 5555 and primed per BAC (BR-127 primer). A 1 inch separator film was installed along one edge of the panels in the bondline for subsequent crack initiation purposes. The test panels were bagged per the BAC processing specification then cured with portable processing equipment using the following cure cycle: 25 inches Hg vacuum pressure applied, 5 F/minute heat-up rate to 250±10 F, a 250 F hold for 90±15 minutes, then ramp down to room temperature. Five individual test coupons were then sectioned from the panels and a tapered wedge carefully inserted into the bond surface at the separator film and a inch crack initiated at the coupon end. Two coupons were placed into a salt spray per BSS 7249 (ASTM B117) and three coupons were left at room temperature. The resulting wedge crack growth data is shown in Table D

40 BAC 5514 calls for the specimens to be exposed to 120±5 F/95% relative humidity for minutes, then measured for crack growth. The BAC 5514 requirement is a 0.3 inch maximum allowable crack growth. The test method used here differs slightly from the method of BAC Nevertheless, we can see that no crack growth occurred in any specimen and the data validates the quality of the PACS anodize surface. Table Wedge Crack Growth, PACS Surface Treatment Validation Coupon Environ- Initial 1-Hour 24-Hour 120-Hour I.D. ment Crack, (in.) Growth, Growth, Growth, (in.) (in.) (in.) W-l-1 Salt Spray No change No change No change W-l-2 Salt Spray No change No change No change W-l-3 Lab Air No change No change No change W-l-4 Lab Air No change No change No change W-l-5 Lab Air No change No change No change Flatwise Tension Strength Flatwise tension coupons were fabricated to provide data for the structural analysis. The flatwise tension coupons were fabricated in two sets; one set with adhesive only and one set with 2 plies of boron/epoxy sandwiched between film adhesive at the bondline. The \ flatwise tension strength coupons were fabricated from 2 x 2 inch aluminum blocks that were phosphoric acid anodized and bonded with BMS Type II Grade 10 (FM-73) adhesive in an oven using the following cure cycle: 25 inches Hg vacuum pressure applied, 5 F/minute heat-up rate to 250±10 F, a 250 F hold for 90±15 minutes, then ramp down to room temperature. The specimens were tested at 0.05 in./min. head speed. No specimens were able to be taken to failure before the pin in the loading heads of the test machine broke. The minimum load at which a pin broke was approximately 16,600 pounds making the specimens capable of at least 4150 psi. Since this test is not typically performed for material characterizations no data was available for comparison. D

41 2.1.4 Conclusions The boron/epoxy material met AMS 3867/4A specification requirements when tested to the AMS specification procedures. The material has similar physical properties as other 250 F curing composite material systems; namely BMS materials. These evaluations indicated that the boron/epoxy behaves similarly to the BMS graphite/epoxy material and, therefore, procedures to install graphite/epoxy materials have a high probability of applying to the boron/epoxy. Typical graphite/epoxy bonded repair procedures were selected as a baseline for the boron/epoxy installation specification based on this reasoning. 2.2 Process Sensitivity Studies Process sensitivity studies were initiated to understand parameters that the doubler application procedures must remain within to develop the design properties. The purpose of exploring the processing window is to develop a robust procedure that is compatible with the austere environments and time constraints typically present in field level support. As was previously stated in Section 2.1.3, the historical difficulties with the application and effectiveness of bonded boron/epoxy doublers has primarily been attributed to the quality of the adhesive bond between the substrate and the doubler and few, if any, problems with the boron/epoxy itself. Therefore the primary factor in evaluating process sensitivities was the performance of the surface treatment and adhesive that provide the bond strength. Primary evaluations included cure temperature and pressure variations, anodize temperature variations, primer cure conditions, and heat-up rate sensitivities Cure Temperature Cure temperature is an issue for the repairs for 3 primary reasons; (1) the aluminum substrate cannot be raised above 260 F without endangering the heat treat, (2) cures over honeycomb structure in excess of the boiling point of water (212 F at 1 atmosphere) will cause the face skins over core with entrapped water to literally blow off due to the generated pressures in the panel, and (3) insufficient heating will not cure the composite materials and preclude development of complete structural capability of the doubler. In reality the cure is controlled by total energy input to the reactive materials so the time at temperature is also important. To isolate the time and temperature variables all other cure variables were held constant at 26 inches Hg vacuum pressure, 5 F/min. heat-up rate to cure temperature, and 5 F/min. maximum cool down to room temperature. The following cure time and temperature combinations were evaluated: D

42 225 F/90 minutes, 225 F/180 minutes, and 200 F/90 minutes. Cure temperatures were held within ±10 F and cure times were within ±5 minutes. Single overlap shear coupons were used to measure the cure effectiveness for each set of conditions. Room temperature shear strengths are summarized in Table Table Overlap Shear Strengths for Alternate Cure Time/Temperatures, RT Cure 225 F/ 90 Min. Coupon Length, Width, Ult. Stress Ult. Load, I.D. (in.) (in.) (lb.) (psi) Avg. Stress,(psi) F/.13HR D

43 180 Min..13HR HR HR HR F/ 90 Min D

44 The baseline average shear strength for coupons cured at 250 F/90 minutes was 4233 psi as reported in Section The 225 F cures for either 90 or 180 minutes provide equivalent shear strengths to the baseline within the accuracy of this low data population. Clearly, the 200 F/90 minute cure does not allow the adhesive to develop adequate properties Cure Pressure Cure pressures for composite materials processed using portable equipment has typically been at full vacuum pressure, 25 inches Hg or higher. Recent data released by the Materials Directorate at Wright Laboratory indicated improved performance of adhesives cured at reduced vacuum pressures. The vacuum pressure variable was therefore selected for evaluation on this program. To isolate the pressure variable all other cure variables were held constant (except as noted) at 5 F/min. heat-up rate to 225 F, 90 minute hold at 225±10 F, and 5 F/min. maximum cool down to room temperature. All aluminum substrates were anodized and primed simultaneously and the primer was airdried. The following vacuum pressures were evaluated: 15 inches Hg, 10 inches Hg, 5 inches Hg, and No vacuum, 4 psi dead weight. Cure pressures were held within ±2 inches Hg and cure times were within ±5 minutes. Single overlap shear coupons were used to measure the cure effectiveness for each set of conditions. Room temperature shear strengths are summarized in Table D

45 Table Overlap Shear Strengths for Alternate Cure Pressures, RT Cure Pressure 5 inches Hg Coupon I.D. Overlap Length, Overlap Width, Ultimate Load, (lb.) Ultimate Strength, (in.) (in.) (psi) Average Strength, (psi) D

46 Table Overlap Shear Strengths for Alternate Cure Pressures, RT (cont.) Cure Coupon Overlap Overlap Ultimate Ultimate Average Pressure 10 inches Hg I.D. Length, Width, Load, Strength, (in.) (in.) (lb.) (psi) Strength, (psi) inches Hg D

47 K psi Dead Wt. 101b-l lb lb lb lb inches Hg/ 1.15HG D

48 250 F 1.15HG HG HG HG The baseline average shear strength for coupons cured at 250 F and 26 inches Hg vacuum pressure was 4233 psi. The average shear strength for the data set of coupons cured for 90 minutes at 225 F and 26 inches Hg vacuum pressure was 4095 psi. All data developed here indicates that reduced vacuum or positive pressure provides significantly improved overlap shear strengths for the adhesive. In fact, all data easily meets the product acceptance test shear strength requirements for BMS Type II Grade 10 adhesive, when the material is cured using autoclave processing techniques. Data generated by Textron and not reported herein (reference Textron memo Q100-TS dated 12/21/94 originated by T. Shahood) indicates that 0 tensile strengths for coupons cured under 28 inches Hg and 15 inches Hg vacuum pressure are equivalent. Based on these observations the baseline cure time, temperature, and pressure parameters were adjusted to 90 minutes, 225 F, and 15 inches Hg, respectively Heat-up Rate Sensitivity Heat-up rates for composite materials processed using portable equipment has typically been at 5 F/minute. The heat-up rate may vary widely in actual practice at an aircraft maintenance facility and the effects of that variation must be understood. To isolate the heat-up rate variable all other cure variables were held constant at 26 inches Hg vacuum pressure, 90 minute hold at 250±10 F, and 5 F/min. maximum cool down to D

49 room temperature. All aluminum substrates were anodized and primed simultaneously. The following heat-up rates were evaluated: l F/minute, and 10 o F/minute. Cure pressures were held within ±2 inches Hg and cure times were within ±5 minutes. Single overlap shear coupons were used to measure the cure effectiveness for each set of conditions. Room temperature shear strengths are summarized in Table The baseline average shear strength for coupons cured at 250 F, 26 inches Hg vacuum pressure, and 5 F/minute heat-up rate was 4233 psi as reported in Section The data for the lower l F/minute heat-up rate was clearly better than the baseline and 10 F/minute heat-up. The improved performance is probably due to a more uniform adhesive layer that was able to cure with fewer voids in the bondline due to a slower rate of volatile release. While faster cures are more desirable in a support environment, the higher heat-up rates provide similar properties to the baseline 5 F/minute rate. The baseline was not changed to require a slower heat-up rate because the practical considerations take precedence over the improved properties. Table Overlap Shear Strengths for Alternate Heat-up Rates, RT Heat-up Rate Coupon I.D. Length, (in.) Width, (in.) Ult. Load, Ult.Stress, (psi) Avg Stress (psi) (lb.) l F/min. 1.1HR HR D

50 1.1HR HR HR F/min. 1.1R R R R / R-5 / D

51 2.2.4 Anodize Parameters The baseline bond strength properties were developed using aluminum coupons that had been phosphoric acid anodized at conditions of 80 F maximum. Emerging data reported by the University of Dayton Research Institute under contract F C-5643 to Wright Laboratory Materials Directorate indicates that the phosphoric acid anodize process has unacceptable performance when performed at 90 F or hotter. Aluminum wedge crack growth coupons were fabricated in accordance with the configuration requirements of B AC One face sheet each was anodized using the PACS equipment per the D process specification at 90 F and 100 F. The panels and anodize solutions were both held at the test temperatures throughout the anodize process. The other face sheet was anodized using the PACS equipment at room temperature. The coupons were then further processed simultaneously and evaluated for the wedge crack growth at room temperature. The coupons were cured using 15 inches Hg vacuum pressure, 5 F/minute heat-up rate to 225±10 F, 90 minute hold at 225 F, and the cool down to room temperature at 5 F/minute maximum. Wedge crack growth results for specimens evaluated at room temperature are shown in Table Table Wedge Crack Growth, Elevated Anodize Temperatures D

52 Anodize Temp. 90 F Coupon I.D. 24-hour Growth, (in.) 72-hour Growth, (in.) 168-hour Growth, (in.) 90-1 No change No change No change No change No change 90-3 No change No change No change 90-4 No change No change No change 90-5 No change No change No change 100 F No change No change No change / No change No change.098 D

53 The baseline wedge crack growth results of Table indicated no change in crack length throughout the evaluation period. The significance of wedge crack growth under constant static load is indicative of the potential for more catastrophic growth (bond failure) in dynamic load applications of real structures. Clearly the 100 F data indicates a significant loss of surface treatment quality and the 90 F data indicates a potential problem as evidenced by the crack growth in one coupon after 1 week. Furthermore, inspection of the failed coupons always showed an adhesive failure to the surface having the elevated temperature anodize. The data indicates that the phosphoric acid anodize procedures should not be performed above 90 F Primer Cure Conditions The BMS 5-89 adhesive primer is typically processed per BAC which calls out an elevated temperature cure of the primer before the adhesive application and bonding. Maintenance time constraints make the added time for normal primer cure less desirable. The effects of flash cures or no primer cure was evaluated. The intent of the flash cure was to use a heat gun with approximately 500 F nozzle exit temperature and heat the primer so fast that it would bubble or blister during cure and simulate a severely overheated primed surface. Overlap shear strength coupons were fabricated to evaluate primer applied to anodized panels and air-dried. The coupons were cured using 26 inches Hg vacuum pressure, 5 F/minute heat-up rate to 250±10 F, 90 minute hold at 250 F, the cool down to room temperature at 5 F/minute maximum. Overlap shear strength results are shown in Table Table Overlap Shear Strength, Alternate Primer Cures, RT D

54 Primer Cure Air Dry Coupon Length, Width, Ult. Load Ult. Stress I.D. (in.) (in.) (lb.) (psi) 1AD Ave Stress (psi) AD AD AD AD-5/ Flash Cure Heat Heat Heat D

55 Heat Heat The baseline average shear strength for coupons cured at 250 F was 4233 psi as reported in Section The air-dried primer coupons provided similar strengths. The primer did not show any visible blistering or bubbles after the flash cure with the heat guns. The flash cure study was for knowledge of the cure sensitivity only. Data generated subsequent to these tests by the Wright Laboratories Materials Laboratory indicate that the air-dried primer may have some problems with bond performance, particularly in wedge crack-growth properties. As a consequence, the doubler installation specification requires a heat cure after primer application and the flash cure results indicate that the heat cure is a robust operation Fluid Exposure of Cured Coupons The boron/epoxy was evaluated for sensitivities to common aircraft maintenance fluids present in service conditions. Cured boron/epoxy 0 compression strength specimens were fabricated from a laminate cured using the evolved baseline cure cycle: 15 inches Hg vacuum pressure, 5 F/minute heat-up rate to 225±10 F, 90 minute hold at 225 F, then cool down to room temperature at 5 F/minute maximum. The specimens were exposed to the following fluids and conditions: Skydrol hydraulic fluid, 30 days at 200 F, MIL-H hydraulic fluid, 30 days at 200 F, MIL-T-5624 (JIM) fuel, 30 days at RT, MEK solvent, 30 days at RT, and MTL-A-8243 anti-icing fluid, 30 days at room temperature. D

56 Specimens were tested at room temperature and the results obtained are shown in Table Table Compression Strength, Fluid Exposure, RT Fluid Coupon Width, (in.) Thickness, (in.) Ult. Load (lb.) Ult. Avg.Stress (ksi) I.D. Stress Control (ksi) MJL-A D

57 Table Compression Strength, Fluid Exposure, RT (cont.) Coupon Width, Thickness, Ult. Load Ult. Stress Avg.Stress Fluid I.D. (in.) (in.) (lb.) (ksi) (ksi) MEK JP D

58 lo MIL-H D

59 Skydrol The 0 compression data indicates that the boron/epoxy coupon strengths are not impacted by exposure to the fluids under the test conditions Conclusions The doubler installation cure conditions were actually improved as a result of the process sensitivity studies and resulted in selecting a final cure cycle of l-5 F/minute heat-up rate to 225±10 F, hold at temperature for minutes, and cool down at 5 F/minute maximum to below 140 F while maintaining a vacuum pressure of 15±2 inches Hg. The boron/epoxy has excellent fluid resistance properties. The process seems fairly robust as long as the cure temperature meets minimum requirements. The process D

60 sensitivity evaluations provide an excellent basis for developing the doubler installation procedure. 2.3 Installation Process Specification The installation process specification, Boeing document number D , is the primary deliverable of the characterization efforts. The specification provides the necessary guidelines to apply high quality boron/epoxy doublers to aluminum structure. As with most specifications some unavoidable situations will be encountered from inservice environments that will require departure from the specification controls. Additional guidance can be obtained from other existing Boeing aircraft support manuals Specification Development The Boeing D "Structural Repair Bonding" document formed the basis for the doubler installation process specification D which appears as Appendix A of this report. From the basic existing procedures, the acquired knowledge of the boron/epoxy material along with investigation of the primary variables expected to be encountered were combined to provide the necessary application controls. The investigation of the variables allowed the baseline cure process to be reduced 25 F in nominal temperature (from 250 F down to 225 F) and vacuum pressure significantly reduced (from 26 inches Hg down to 15 inches Hg) to provide improved structural properties. In addition, lessons learned during structural validation coupon testing allowed further improvements in the specification controls Specification Validation The controls of document D were employed to fabricate the majority of the structural test coupons. In addition, an early version of the specification was used in the application of 25 doublers on 2 operational Federal Express (Fed Ex) 747 freighter aircraft in early 1993 using Fed Ex technicians with Boeing engineering witnesses. Service history of these doublers have been reported in numerous proceedings of which the following 2 references are provided: D

61 "Status of Bonded Boron/Epoxy Doublers for Military and Commercial Metallic Aircraft Structures," E. Bruce Belason, proceedings from SAE Airframe Finishing, Maintenance and Repair Conference and Exposition, March 13-16,1995, "Evaluation of Bonded Boron/Epoxy Doublers for Large Commercial Transport Aircraft," Bruce Belason et al, proceedings of the 5 International Conference on Structural Airworthiness of New and Aging Aircraft, June 16-18, Fabrication of Field Repair Test Specimens The field repair and static fatigue test coupons (see Section 4.0) were fabricated according to the procedure shown in Figure The specimen fabrication procedure had to be amended after the first set of specimens were fabricated, tested, and analyzed. The failure analysis after testing of the first group of test specimens discovered that a large portion of the coupons had cracks that propagated from the stop-drilled hole periphery at burrs or other hole imperfections that were not visible to the unaided eye. The hole inspection operation number 080 was added to the fabrication procedure for all other subsequent test coupons so that sources of variation would be minimized. The microscopic inspection operation is not practical for actual field repairs and reinforces the need to use good quality drills and bits, reaming when necessary, and properly skilled technicians in actual practice. Stop-drilled holes were not intentionally filled during specimen fabrication though some adhesive flash generated during the doubler cure operations was present in all holes. Coupon anodizing was performed using a tank process per B AC 5555 rather than the PACS system for greater efficiency since the two methods have been shown in previous Boeing characterizations to provide equivalent surface features and adhesive bond strengths. All cures were performed using a portable repair controller (cure equipment) and heat blankets. Op. No. Operation 010 Draw inch thick 7075-T6 aluminum sheet per QQ-A-250 from stores. 020 Layout coupon cutting pattern on the sheet per the appropriate engineering drawing (SK-BREP-001, SK-BREP-002, etc.). Ensure proper grain D

62 direction for the blanks. 030 Cut coupon blanks to drawing dimensions. 040 Layout 0.25 inch diameter hole on each coupon per the appropriate engineering drawing (SK-BREP-001, SK-BREP-002, etc.). 050 Drill 0.25 inch hole in each coupon at prescribed location. 060 Break all sharp edges on panel periphery per BAC Using a jeweler's saw install a inch maximum width saw cut perpendicular to the longitudinal edge and traversing from the near edge to the 0.25 inch diameter stop-drill hole. Terminate the saw cut in the center of the hole. 080 Visually and microscopically inspect the periphery of each hole for knicks, burrs, or other imperfections in the metal. Reject all blanks with noted imperfections. 090 Phosphoric acid anodize each coupon blank per BAC Verify anodize quality with polariscope. 100 Draw BMS 5-89 primer from storage and allow to warm to room temperature until no condensation exists on the container. 110 Apply BMS 5-89 primer to the anodized aluminum blanks per BAC Air dry the primer for 2 hours minimum. Do not heat cure the primer after air drying. 120 Place the blanks in pairs with primed surfaces together. Wrap the paired blanks in clean Kraft paper. D

63 Figure Fabrication Procedure for Field Repair Test Specimens p. No. Operation 130 Remove the 5521/4 boron/epoxy unidirectional tape from the freezer. Allow the material to warm to room temperature until no condensation is present prior to removing the protective wrapping. 140 Solvent clean the aluminum cutting templates per BAC 5750 with methyl ethyl ketone (MEK). 150 Cut the plies required for each doubler using the appropriate templates. Ensure proper fiber alignment in each ply. Change utility knife blades often to avoid damaging the material. 160 Lay up each set of plies per the appropriate engineering drawing (SK-BREP- 001, SK-BREP-002, etc.). Debulk the ply stack under 22 inches Hg minimum vacuum pressure for 30 minutes minimum. 170 Remove the compacted ply stack from the vacuum debulk and apply 1 ply of BMS Type II film adhesive to the doubler bond surface. Debulk the doubler lay-up under 22 inches Hg minimum vacuum pressure for 10 minutes minimum. 180 Remove the doubler lay-up from the vacuum debulk. Apply 1 ply of style nylon peel ply to the exterior surface of the doubler (the surface without adhesive). Sweep out all wrinkles in the peel ply. 190 Draw a primed aluminum blank. Apply the doubler to the blank per the appropriate engineering drawing (SK-BREP-001, SK-BREP-002, etc.). 200 Place completed lay-ups in the hot bond repair console vacuum bag per Figure 9 of the D process specification. Place a minimum of 3 monitoring thermocouples on the coupons adjacent to but not on the boron/epoxy doublers. D

64 Figure Fabrication Procedure for Field Repair Test Specimens (cont.) Op. No. Operation 210 Program the controller for the following cure cycle: 15 inches Hg vacuum pressure (±2 inches) l-5 F/minute heat-up rate Hold at F for minutes Cool-down at 5 F/minute maximum Cure complete when below 140 F 220 Cure the coupons. Monitor vacuum pressure and thermocouple readings. 230 Debag the coupons after cure. Remove the peel ply and break all sharp edges and protruding fibers. 240 Visually inspect the coupons for adhesive and resin flow and general appearance. Reject all coupons that have unusual appearance or contain doublers that have shifted more than 1/4 inch from their intended position. 250 Wrap completed coupons in protective packaging. 260 Deliver to Paul Rutherford, Figure Fabrication Procedure for Field Repair Test Specimens (cont.) 3.0 Nondestructive Inspection (NDI) Two different nondestructive inspection (NDI) techniques are used to inspect the boron/epoxy doubler, adhesive, and aluminum substructure: ultrasonic resonance and eddy current. 3.1 Ultrasonic Resonance D

65 NDI requirements for the inspection of anomalies in or under the doubler were given by Peter Murphy of BCAG Quality Assurance Research & Development (QAR&D). Textron developed the NDI ultrasonic procedure required to detect: (1) disbond of the adhesive layer between the boron/epoxy doubler and metallic parent structure; and (2) the delamination of the composite boron/epoxy doubler between the ply layers. NDI reference standards are required for instrument calibration to detect the anomalies the inspector is inspecting for. Figure is a sketch of the reference standard designed which would test the delectability of 1 inch diameter and 0.5 inch diameter disbond and delamination anomalies. Table shows ten reference standards with a different number of boron/epoxy doubler plies on different parent material thicknesses. Typically, commercial detectable delamination standards are 1 inch diameter voids. However, it is envisioned that some doubler installations would be small enough that it would be desirable to detect an anomaly smaller than 1 inch diameter. Part of the reason a 1 inch diameter void is the detectable flaw size is because an inspector is typically inspecting a large region (an example would be bonded body skin lap slices). To simulate the disbond region it was recommended by Boeing QAR&D to try two different methods on a test fabrication standard. The two methods are: (1) use back-toback FEP 2-3 mil solid Teflon shims, and (2) apply free-coat 700 (room temperature cure) to a 1-mil brass shim and apply it in the bond region. For the edge delamination areas, the brass shims can be pulled out. The reference standard was successfully fabricated, and the ultrasonic procedure was developed at Textron. Textron and Federal Express personnel used the procedure to inspect 25 boron/epoxy doublers installed by Federal Express onto two operational Federal Express 747s in early D

66 Figure Boron/Epoxy Nondestructive Inspection Reference Standard for Disbond and Delamination Evaluation Table Boron/Epoxy Reference Standards Required for Disbond of the Adhesive and Delamination of the Composite Ref. Std# Alum. Thick Layup # Plys Doubler Thick %Br/Ep Depth Depth #1 # (02)s % 2-to-3 3-to (02)s % 2-to-3 3-to (03)s % 2-to-3 4-to (03)s % 3-to-4 5-to (04)s % 3-to-4 5-to (+45, % 2-to-3 5-to-6 45,02)s (04)s % 3-to-4 6-to (05)s % 3-to-4 7-to (05)s % 4-to-5 8-to (06)s % 4-to-5 10-to Eddy Current Textron brought to Boeing the boron/epoxy step wedge specimen Textron had been using to evaluate different eddy current responses for different size and types of flaws under the doubler. Eddy current indications from different flaw sizes and depths were observed using a 0.2 inch diameter (5 khz) spot probe and an encircling probe (2-5 khz). It was D

67 obvious that certain crack types and sizes under some boron/epoxy thicknesses can be found reliably. We discussed the design of a reference standard to aid in developing the eddy current NDI procedure for the detection of crack growth propagating from a stop drill, rivet, or freeze plug. Figure shows the design of the reference standard we discussed. Figure Boron/Epoxy Reference Standard Eddy Current NDI for Crack Growth Detection Below the Doubler 4.0 Structural Analysis and Performance Tests A significant portion of the data gathered on this contact was empirical from testing accomplished at the Intec test facilities in Bothell, Washington. However, analysis tools such as INCAP, COSMOS, NIKE/3D, and ANS YS were used to aid in the design of the test specimen geometries and to help understand fabrication-induced stresses. The testing D

68 is separated into two subjects: boron/epoxy material property testing and field repair testing. Limited material property testing was performed since property data existed for boron/epoxy tape using a 350 F cure temperature. The bulk of the testing performed was on the field repair specimen which was a boron/epoxy doubler bonded to a damaged 4x16 inch thin aluminum sheet. 4.1 Material Property Tension Tests on Vacuum-Bag Cured Boron/Epoxy Boron fiber material used in other aerospace vehicles utilized a 350 F cure resin tape cured in an autoclave. Material property data for this tape existed in the literature. However, for on-aircraft co-cured repairs the 250 F epoxy tape is of interest because temperatures above 260 F can damage the heat treat of the aluminum structure. Little material property data was published for the boron fiber with the 250 F epoxy tape cured with a heat blanket and a vacuum bag. Textron published information for the 250 F epoxy system 1 shows little difference with the 350 F epoxy system, see figure 4.1-1, when cured in an autoclave. Since the boron/epoxy doubler is in tension during normal application, static tension testing was the material property test proposed on the phase I test plan (see Section 1.2) and performed on this task. Figure shows the test matrices for the material property tests accomplished. Two specimen configurations with different fiber lay-ups, unidirectional and bow tie (0 ±45 ), were exposed to four different environmental conditions: -75 F, room temperature, 180 F/wet, and solvent soak. Test results in several environments for the 350 F epoxy system were published in the NASA/DOD Composite Design Guide and are shown in figure This material is designated #5521 by Textron. D

69 Source of Property Values 1 Br/Ep, B(4) fiber and N5505 matrix (350 F). From Tsai, Composite Design 2 Br/Ep, NARMCO 5505 matrix (350 F cure). From MIL-HDBK-17A, Average Prop.75 F 3 Br/Ep Prepreg Tape-5505 (350 F cure), from Textron Vendor Handout, RT 4 Br/Ep Prepreg Tape (250 F cure), from Textron Vendor Handout, RT Source of Property Values Ply Engineering Constants E11 (msi) E22 (msi) 2.68 E33 (msi) G12 (msi) 0.81 G23 G13 (msi) (msi) u! u.23 u!3 Other Ply Data v/f 0.5 P (Ib/in3) t (mils) D

70 Figure Material Property Values for Different Boron/Epoxy Tapes Source of Property Values Max. Strength and Strain ol (ksi) Prop.Lmt. (ksi) ale (ksi) (ksi) a2c (ksi) 29.3 a3 (ksi) a3c (ksi) Tl2 (ksi) T23 (ksi) Tl3 (ksi) el (E-3 in/in) elc (E-3 in/in) e2 (E-3 in/in) E2C (E-3 in/in) e3 (E-3 in/in) e3c (E-3 in/in) el2 (E-3 in/in) 11.9 e23 (E-3 in/in) 813 (E-3 in/in) Figure Material Property Values for Different Boron/Epoxy Tapes (cont.) MATERIALS: Doubler Material = Boron/Epoxy 5.0 mil fiber Tape VARIABLES: Boron/Epoxy = Environment = Batch Number Boeing Test Standard Tensile Testing of Advanced Composites Extensiometer on all specimens [0/45/0/-45/0]s Et = 966 kips/inch Batch Number D

71 Tension Testing Instrumentation Br/Ep Ply Layup [0/45/0/-45/0]s -75 F, Lab Air [0]8 Room Temp, Lab Air 180 F/Wet Solvent Exposure Test Temp/Hum Number 1 Number 2 Number 3-75 F, Lab Air Room Temp, Lab Air F,Wet Solvent Exposure Br/Ep Ply Layup [0]8 Et=1183 Batch kips/inch Number Test Temp/Hum Number 1 Number 2 Number 3-75 F, Lab Air Room Temp, Lab Air F,Wet Solvent Exposure Total Samples for Material Testing = 120 Figure Material Property Tension Test Matrices D

72 Materials Properties from NASA/DOD Composite Design Guide Ply Engineering Constants A-Value Allow. Average Prop. Average Prop. Average F E11 (ma) E22 (ma) G12 (ma) u Other Ply Data v/f r (Ib/in3) t (mils) Max. Strength and Strain a1 (ksi) a1c (ksi) CT2 (ksi) a2c (ksi) T12 (ksi) e1 (E-3 in/in) E2 (E-3 in/in) Hygrothermai Expansion coefficient a1 (E-6 in/in/ F) a2 (E-6 in/in/ F) Figure Material Property Values Published in the NASA/DOD Composite Design Guide D

73 4.1.1 Environmental Testing The -75 F air conditioning was done prior to testing in the frame. The specimens were tested in a frame with an integral environmental chamber which used liquid nitrogen (LN 2 ) to cool it. Thermocouples were mounted on the specimens. When they achieved - 75 F the test was performed. The room temperature air condition is slightly variable at the Intec test facilities, however, conditions are normally 70 F with low humidity. The 180 F/wet specimens were conditioned in an environmental chamber at 160 F and 85% RH (standard conditioning and testing) at Intec. This conditioning allows the specimens to absorb water. The moisture intake on these specimens was monitored using travelers (a specimen which is used only for moisture content predictions). After conditioning, the specimens were tested at 180 F/dry. Appendix B contains the data on these travelers. The specimens were tested when the travelers indicated that they had stabilized at an equilibrium moisture content. The solvent exposure specimens were soaked in Skydrol hydraulic fluid for 7- days at 120 F at Boeing Material Technology Lab. After the week soak, the specimens were plastic wrapped and delivered to Intec to be tested within 8 hours after removal from the Skydrol exposure Unidirectional Tension Specimen Tests Sixty unidirectional ply test specimens were fabricated from three 8 ply panels. The specimen configuration employed bonded on tabs as shown in figure The specimens were tested using ASTM D3039 and BSS 7320 (Tensile Testing of Advanced Composites) as guides. A photograph of the test setup is shown in figure The loading rate used was 0.05 in/min. Strain was recorded digitally until failure using an extensometer, with a 1 inch gage length, mounted on the face of the coupon at midsection. Low failure stress levels (approximately 94 ksi versus the 220 ksi Textron published value for autoclave cured material) were occurring at the tabs on these specimens. These values are summarized in Appendix C. Figure shows this tab failure. On closer inspection, insufficient taper on the tabs was identified as a probable source of the problem, see Section Testing with the short tab taper specimens stopped until the problem was determined. To test this hypothesis, a quick test was D

74 performed at Intec. The tabs were cut off a spare specimen and tested. The grip pressure was minimized to avoid damaging the surface fibers. The failure stress level increased to 166 ksi from 94 ksi. The unidirectional specimen tab taper was investigated and is reported in Section However this investigation still resulted in low values (approx. 160 ksi). At this point all 0 tension test with tabs were terminated. Textron fabricated a new set of 0 tension specimens and performed testing using a unique test procedure they developed which has straight sided tabless specimens. During this test, a steel mesh is wrapped around the grips to reduce stress concentrations 2. The results of Textron's tests are published in Textron Memo No. Q100-TS Figure Unidirectional Ply Tensile Test Specimen Figure Photograph of Test Setup for Unidirectional Tension Tests 2 This procedure is identified by Textron as Quality Assurance Test Procedure #5171 D

75 Figure Photograph of Tab Failure on Unidirectional Tension Specimen Bow-Tie (0 +45 ) Tensile Specimen Tests Sixty cross ply [0/45/0/-45/0/0/-45/0/45/0] test specimens were fabricated from three panels. The specimen configuration is shown in figure Section contains the laminate analysis which determined the approximate stiffness off the specimen. The specimens were tested using ASTM D3039 and BSS 7320 (Tensile Testing of Advanced Composites) as guides. A photograph of the test setup is shown in figure The loading rate used was 0.05 in/min. Strain was recorded digitally until failure using an extensometer, with a 1 inch gage length, mounted on the face of the coupon at mid-section. These cross-ply specimens yielded failure stress around 120 ksi. The laminate analysis shown in Section shows that the 0-degree plies were being loaded to approximately 218 ksi at the failure load, which is close to expected values. Figure Bow Tie Tensile Test Specimen A summary of the average ultimate tensile stress and tensile modulus from the bowtie tension tests are shown in figure and respectively. Most of these specimens failed in the gage section. A photograph of a typical failed specimen of this type is shown in figure D

76 Appendix C contains the data from the bow tie tension tests and the statistical analysis for comparisons of the different environments (treatments), and to see if there is a statistically significant difference between the three different panels fabricated. Four dependent response variables were measured from the test: ultimate stress, ultimate axial strain, modulus of elasticity (MOE) using least squared fit (LSF), and chord MOE. The independent variables were environment and the panels made from different boron/epoxy manufacturing batches. From this data it can be seen the bow tie specimens tested from Panel f (A130 batch) shows significantly lower strength and modulus of elasticity (MOE) values. After further failure investigation by Textron, it was concluded that the Panel f specimens were fabricated with the scrim-cloth on the specimen which created a thicker than normal specimen and yielded lower values. The hot/wet and solvent test environment show the lowest strength and MOE values, but they are not significantly different from each other. Figure Photograph of Test Setup for Bow Tie Tension Tests D

77 Figure Summary of the Material Property Results for the Bow Tie Tension Tests Tensile Strength Property Results Figure Summary of the Material Property Results for the Bow Tie Tension Tests Tensile Modulus Figure Photograph of Typical Failure in Gage Section of a Bow Tie Tension Specimen D

78 4.2 Stress Analysis Stress analysis performed on this contract was in support of the material property and field repair test programs. Four different analysis software packages were utilized: INCAP, COSMOS, NIKE, and ANSYS. INCAP is a composite laminate analysis program which applies laminate plate theory to determine the stiffness matrices for the laminate. This was utilized when sizing the material property specimens and different boron/epoxy doublers used in the field repair testing. COSMOS is a 2D linear finite element program, which is a commercial produced software, and was run on a Macintosh. It was initially used to get an indication of the internal stress due to the cure of the composite doubler on an aluminum base. NIKE is a 3D nonlinear finite element code which is supported by Lawrence Livermore National Laboratory. This code was used to size different baseline specimen geometries considered for testing. Since it is a 3D code, it accounted for the crack size and the edge margin of the doubler perpendicular to loading. ANSYS is another finite element commercial product used to further understand the thermal stresses produced during fabrication and how tension mechanical loading opposes this thermal loading. Of concern early in the analysis task was the linear material property assumption made in the finite analysis methods. The aluminum and the boron/epoxy composite behave linearly with slight changes over a large temperature range. However, the adhesive used to transfer load from the aluminum structure to the boron/epoxy doubler is very nonlinear over large strains and temperature ranges, see figure Also of concern is that the Boeing BMS film adhesive specification has a number of different vendors that meet that specification. However, from an analysis viewpoint each film adhesive from each vendor has different material properties used in an analysis. Figure shows three different BMS film adhesives with a few of their material properties. Because of the large variations in the properties, it was decided to use only American Cyanamid FM-73. This is the adhesive procured for this test program and the adhesive used for the tests accomplished in Section 2 of this report. D

79 Figure Shear Stress vs. Shear Strain for American Cyanamid FM73 Film Adhesive Vendor Type E (psi) G (psi) H a Hysol EA , Amcy FM73 310, , M AF ,000 60, Figure BMS Film Adhesive Material Properties Reported from Vendors The following gives a brief description of the stresses in the adhesive of a bonded joint. Figure shows generally the different adhesive stresses (face or in-plane, shear, and transverse or peel) which occur during a single lap joint in tension. The magnitude of these values are a function of the stiffness (material properties and geometry) of the adhesive and the adherents. There are different approximation analysis methods to determine the stress distribution in the adhesive during loading until failure. Figure show the adhesive normalized shear and peel stress distribution at failure for a single shear lap joint analyzed with four different methods: (1) LE FEM, is linear elastic finite element method; (2) NL FEM, is nonlinear elastic finite element method; (3) Kreiger, is Dr. Kreiger's analysis method; and (4) Avg. Shear, is the load averaged over the area of the bond. Typically adhesive shear stress data is presented in the fourth method. The finite element analysis contained in this section uses the linear elastic and nonlinear methods. However, the nonlinear method accounts for geometric nonlinearities caused from the deformation of the finite element mesh during loading and not material nonlinearities with the film adhesive used over both different strains and different temperatures. D

80 Figure Generalized Single Shear Lap Joint Stress Figure Normalized Shear Stress vs. Distance Along a Single Shear Lap Joint in Tension INCAP Laminate Analysis INCAP is a composite laminate analysis program which applies laminate plate theory to determine the stiffness matrix for the laminate. This was utilized when sizing, the material property specimens and different boron/epoxy doublers used in the baseline field repair testing. Appendix E contains the analysis performed when sizing the material property bow tie (0 ±45 ) specimens, see Section 4.1, and the baseline doubler geometries tested, see Section COSMOS 2D Finite Element Analysis COSMOS 2D linear finite element program was run on a Macintosh and was used to get an indication of the internal stresses due to the cure of the composite doubler on an the aluminum structure. Appendix F contains a summary of the analysis performed with conclusions. A few of the results found from this analysis are: D

81 Thermal processing stresses the doubler in the opposite direction than during axial tension loading. Therefore, the thermal stresses created from fabrication help to lower the peak stresses in the doubler resulting from tension loading. The decrease in the maximum principal stress in the boron doubler was 30-40% from the axial load case without thermal fabrication stresses, which also results in lower adhesive and aluminum stresses. Also observed were that doublers near hard points such as stringers or frames had increases in adhesive shear stresses by more than 75% and increases in aluminum bending stresses by more than 200%. The stresses in the aluminum in front of the doubler can be high. Lastly, a few analysis runs were used to investigate doubler ply drop-off. It was found that the highest shear and bending stresses occur with the smallest ply dropoff 1:10, which is expected. No maximum ply drop-off can be recommended, however minimum ply drop should be greater than 1:10, say 1:20 or more NIKE 3D Finite Element Analysis NIKE is a 3D nonlinear finite element code which is supported by Lawrence Livermore National Laboratory. This code was used to size different baseline specimen geometries considered for testing. Since it is a 3D code, it accounted for the crack size and the edge margin of the doubler perpendicular to loading. This analysis was used to help determine the baseline doubler geometries for the specimens to be tested. Appendix G contains the data gathered from this analysis. A few interesting results were determined from this analysis: Both cases were studied where the doubler covered an edge crack and a center crack. No significant difference exists in the resulting material stresses. This means that the edge crack test and analysis results may be applied in designing doublers for centrally cracked panels. Peak stresses in the adhesive occur at the doubler leading corner. Shear stresses in this region are 7-8 times higher than stresses away from doubler edges. The maximum peel stress of+1.2 ksi also occurs here, and rapidly drops to a slightly negative value over most of the adhesive. These peak stresses could probably increase with added mesh density, which would more accurately predict the stress gradients. D

82 Stress in the aluminum in front of the doubler is at least 25% higher than the applied axial stress. This peak occurs approximately 0.5 inches in front of a 4 inch long doubler ANSYS 2D Finite Element Analysis ANSYS 2D finite element program was used to further understand changes in the stresses during fabrication due thermal loading and how tension mechanical loading opposes this thermal loading. Appendix H contains a summary of the results from the analysis performed. This stress analysis was requested by Textron to investigate the stresses induced during cure and during a 15 ksi axial loading. The analysis was accomplished for three different boundary conditions (near stringer support, far stringer support, and free edge support) and three different load cases (stress after cure, stress from a 15 ksi axial load, and t two combined). Color stress distributions are shown in Appendix H for each load case with the near stringer boundary condition, which yields the higher stresses. 4.3 Field Repair Tests Field repair testing (also called performance testing) is divided into two separate sections: baseline geometry tests and parametric tests. The baseline geometry tests examined three different configuration variables relating to the design of the boron/epoxy doubler: ply lay-up, ply orientation, and doubler geometry. This testing is labeled as Phase One in the field repair test matrix, Appendix D. It was the intent in this testing to culminate in selection of one design for further evaluation in the parametric testing. This design was called the Baseline Design. The parametric tests were designed to examine different variables on the selected baseline design. Some of the variables examined were: doubler ply reduction, aluminum thickness changes, different crack lengths, environmental exposures, different doubler anomalies, doubler impact events, and disbonds. These tests are labeled as Phases 1.5, 2, 3, 3.5,4, and 5 in the field repair test matrix contained in Appendix D. D

83 4.3.1 Baseline Geometry Tests The baseline geometry tests were separated into three different test methods: static tension tests, tension-tension fatigue, and impact with either a static tension test or a tension-tension fatigue test after the impact event. The baseline geometry field repair specimen was a boron/epoxy doubler bonded onto a 16x4 inch 7075-T inch thick aluminum sheet with a 0.5 inch saw cut centered from the edge of the aluminum with a 0.25 inch diameter hole at the end of the saw cut centered under the doubler. The doubler was co-cured with the FM73 film adhesive at 225 F with 15 inches Hg vacuum (see Section 2.4). Each boron/epoxy doubler configuration was tested over this same aluminum substrate. The number of plies and the ply orientation of each doubler lay-up is a function of the loading on the structure and the stiffness of the structure being repaired. The procedure used while installing "decal" doublers on a Federal Express 747 was the following: match the cross-sectional stiffness, Et, with the surrounding structure, then round up to the nearest whole ply, and add an extra cover/wear ply. This is how the unidirectional boron/epoxy doublers were sized for the field repair specimens. The three variables investigated in the baseline geometry determination tests were: ply layup, ply orientation, and doubler geometry. There were two different ply lay-ups: normal and inverted, see figure The normal lay-up had the area of each ply reducing toward the top of the doubler until the last cover ply which covered the stacked plies and vice-versa for the inverted. Two boron/epoxy ply orientations were investigated: a 0 (i.e. parallel to the load, hence perpendicular to the crack), and (also called "cross-ply lay-up"). The doubler geometry was either a rectangular doubler or a rectangular doubler with cut-off corners, see figure , and was identified on the specimen by a 1, 2, or 3. Cut corners instead of curved corners were employed because of the difficulty cutting curves in the boron/epoxy tape. In figure , the difference between doubler geometry 1 and doubler geometry 2 is the over all size of the doubler. For a detailed description of each specimen, see Appendix I. It contains all the field repair drawings for the specimens tested. The laminate analysis of the cross-ply configurations is contained in Appendix E. Six replicate specimens were fabricated for each of the twelve different doubler configuration (combination of 2 x 2 x 3 variables) giving 72 specimens fabricated to test. Each specimen was through-transmission ultrasonic (TTU) nondestructively inspected (NDI) for delamination in the composite doubler or disbond of the doubler to the aluminum D

84 structure. Appendix J contains the TTU NDI C-scans for some of the field repair specimens inspected. The specimen numbering system reflects the specimen doubler geometry. Figure shows the specimen number for each of the different doubler configurations fabricated to be tested. Figure Lay-up Configurations and Doubler Geometries for Field Repair Static D

85 Specimen # Plys Layup Edge Margin, (inches) * 1A [0]6 L = 2.0,W = 1.5 Normal 1B [0]6 L = 2.0,W=1.5 Inverse 1C [0/45/0/-45/0/-45/0/45/0] L = 2.0,W = 1.5 1D [0/45/0/-45/0/-45/0/45/0] L = 2.0,W = 1.5 Normal Inverse 2A [0]6 L = 3.0, W = 2.5 Normal 2B [0]6 L = 3.0, W = 2.5 Inverse 2C [0/45/0/-45/0/-45/0/45/0] L = 3.0, W = 2.5 2D [0/45/0/-45/0/-45/0/45/0] L = 3.0, W = 2.5 Normal Inverse 3A [0]6 L = 2.0, W=4.0 Normal 3B [0]6 L = 2.0, W=4.0 Inverse 3C [0/45/0/-45/0/-45/0/45/0] L = 2.0, W=4.0 3D [0/45/0/-45/0/-45/0/45/0] L = 2.0, W=4.0 Normal Inverse * see Appendix I for details Figure Numbering System for Fatigue Specimens with Lay-up Configuration and Doubler Geometry Static Tests All field repair static specimens were tested using ASTM D3039 as a guide. The specimen dimension where the 4 inch hydraulic wedges attached to the specimen is shown in figure This is also the wedge location during the fatigue tests discussed in Section The loading rate was 0.05 in/min. Instrumentation for most static tests consisted of an extensometer (0.5 inch gage length) mounted across the stop drill. In addition to this, a strain gage was also mounted in-line with the stop drill, half way to the edge of the doubler. The location of the instrumentation on the static field repair specimens is shown in figure The test setup for a static field repair specimen is shown in figure D

86 Side view of specimen is curved due to thermal stress created during doubler installation. Figure Specimen Dimensions for Field Repair Static and Fatigue Tests D

87 Figure Photograph of Field Repair Static Test Setup Thirteen static tests were conducted. Each specimen configuration was tested once, except configuration ID which was tested twice. The dependent test variables from the static test were: ultimate strength, modulus of elasticity (MOE) from the strain gage, MOE from the extensometer, and global MOE for the entire specimen. Figure shows a representative load vs. strain plot of the static field repair specimen. Each of the twelve different doubler designs restored the flawed 7075-T6 aluminum specimen to its full static ultimate strength. A summary of the ultimate tensile stress and global tensile modulus are included in figure and , respectively. The global tensile modulus 3 is less than that of aluminum alloy which is 10.5 msi. This is important in a structural repair because typical repair doublers are designed to restore the overall MOE repair to that of the original structure. Appendix K contains a table of the baseline field repair static tension test results and the statistical analysis performed on that data. The four dependent response variables analyzed were: ultimate stress, strain gage MOE, extensional MOE at the stop drill, and global MOE. The 12 independent treatment variables were: plies (0 's or 's), lay-up (normal or inverted), and geometry (1, 2, or 3). 3 Use gage length of aluminum substate and boron/epoxy doubler to calculate. 11 inches typically. D

88 The analysis of variance (AOV) tables for the treatments show a significant difference for the different geometries with the ultimate strength and global MOE response variables. Geometry 2 always has the highest mean value for each of the dependent response variables, though it may not always be statistically significant from all other geometries, the trend for Geometry 2 to yield the highest value is consistent. The lay-up independent variable is never statistically significant and therefore has no influence on the static test response variables. The ply orientation independent variable is statistically significant with extensional MOE at the stop drill and global MOE. For extensional MOE at the stop drill, the higher value for the 0 degree lay-up will reduce the stress intensity at the stop drill. For D

89 global MOE the 45 degree plies more closely approaches the undamaged aluminum MOE of 10.5 msi, however, the 0 degree ply doubler is not much lower. Based on the static field repair test results, the independent variables which would be preferred are: Geometry 2, with 0 degree plies, and lay-up does not matter Fatigue Tests All field repair fatigue testing was performed using ASTM D3479 and ASTM E647 as guides. All the baseline fatigue tests were constant amplitude at 5 Hz. Initially the fatigue specimens were instrumented with strain gages in addition to crack detection gages at the edge of the hole. Strain surveys were conducted on the fatigue specimens at specified intervals to see if any mechanical hysteresis had occurred, see figure This practice was discontinued because no mechanical hysteresis was observed. Hence, for the remaining specimens instrumentation was restricted to crack detection gages and crack propagation gages only. Also, the first three fatigue specimens were tested between 3 and 15 ksi (R=0.2) for 300,000 cycles. Since no crack was observed, the maximum stress level was increased to 20 ksi (R=0.15). 20 ksi is also the maximum operating stress for some commercial fuselage skin structures. A crack detection gage was applied to the edge of the simulated stop drill, 0.25 inch diameter hole, see figure The Measurements Group CD-02-20A crack detection gage was used with MBond 200 adhesive. The crack detection gage provided an economical method of indicating the presence of a crack at the stop drill hole. However, sometimes the crack would not be detected until the crack tip was significantly past the gage filament. This was significant in a few cases, once where the crack advanced almost a half an inch before detection. D

90 Figure Photograph of Setup for Field Repair Fatigue Test with Crack Detection Guage The number of cycles to crack detection was recorded and upon crack detection the specimen was removed from the test frame. Crack propagation gages were then installed and crack growth monitored with constant data acquisition. Usually two in-line crack propagation gages were used per specimen, when space permitted, see figure Measurements Group TK-09-CPC DP with 20 resistor strands was mounted on the specimen with Measurements Group MBond 200 adhesive. This pattern of gage was preferred because of the area of coverage and greater uniformity of increases in total resistance with successive strand fractures. The baseline constant amplitude fatigue minimum and maximum stress levels were 3 and 20 ksi. These field repair specimens had residual curvature (see figure ) associated with the residual thermal stresses due to the 225 F cure temperature and the different coefficients of thermal expansion of the boron/epoxy and the aluminum. The selection of 3 ksi at room temperature as the lower limit was to ensure that there would not be any significant stress reversals during constant amplitude fatigue testing. A stress level of 3 ksi seemed to significantly straighten out the inch thick aluminum specimen with a doubler configuration 1A (see figure ) bonded on it. 20 ksi was chosen as the higher limit to equate to maximum fuselage operating stresses. Runout was established as 300,000 cycles (more then 4 aircraft lifetimes) for these fatigue tests. D

91 Residual strength static tests were done on all the specimens that ran out or failed fatigue compliance and remained in one piece. This test was accomplished like the previous static tension tests described in Section Figure Photograph of Setup for Field Repair Fatigue Test with Crack Propagation Gages Appendix K contains a table of the baseline field repair fatigue test results and the statistical analysis performed on that data. The statistical analysis was performed on the two response variables gathered during this test: cycles to crack initiation, or actually crack detection, and total cycles. The crack growth curves were analyzed graphically. From both of the response variables analyzed and shown in Appendix K, it can be concluded that there is no statistically significant difference between each of the different treatments: geometry, plies, or lay-up. However, Geometry 2 always yields a greater number of cycles until crack detection and a greater number of total cycles, which is better performance, even though this is not statistically significant, which means that there is too much variance in the data to say that one treatment is better than another with at least a 95% probability of being correct. For this reason it can be concluded that Geometry 2 would be the desirable doubler configuration, which agrees with the static tests, and D

92 specimen configuration 2A was selected as the baseline specimen design for the parametric testing in Section Appendix K also contains the data and the plots used to analyze the crack growth behavior for the different field repair specimens on which cracks were detected before 300,000 cycles. The crack growth data in Appendix K is presented in two different ways: first, the typical total crack length vs. total cycles plot; second, normalizing the number of cycles to a crack length of 1 inch. Specifically, the number of cycles to grow the crack to 1 inch was subtracted from the total number of cycles for each data point and then plotted. This projects all the data to a common reference, so they can be compared graphically. After observing the data presented in Appendix K the following conclusions were made: all specimen configurations show similar crack growth rates from crack detection to a crack length of about 2 inches (i.e. until the crack reaches the ply drop-off edge of the narrowest doubler); crack growth rates are constant until the crack progresses into the ply drop off edge margin (the point where the doubler starts tapering down): 1.5 inch for geometry 1 and 2.5 inches for geometry 2; inverted versus normal lay-up and 0-degree versus cross plies show no difference on crack growth rates for the specimens on which cracks were detected; once the crack grows beyond the edge of the doubler, the specimen fails within a few more cycles. Also enclosed in Appendix K is the completed static residual strength test results for the specimens which endured fatigue testing for 300,000 cycles without failure. The data shows little difference between different residual strengths, and all are higher than A and B allowables, with an average of 83 ksi Impact Screening Tests Six spare specimens not used in the static and fatigue tests of Section and were used to determine the impact level for tests to be performed in the parametric testing (see Section for parametric impact testing). The impact kinetic energy level used on these spare specimens was 100,300, 500, and 1200 in-lbs. All these specimens D

93 were impacted on a 'Dynatup 8250' drop weight instrumented impactor. Impact coupons were mounted on a 0.75 inch thick aluminum fixture with a 3 x 5 inch cutout window. They were held in place by a second aluminum window and rubber clamps. The impact tup was made of A2 tool steel hardened to 55 RC and was of a 1 inch diameter hemispherical geometry. The location of the impact, and the location of the window relative to the specimen is shown in figure After impact these specimens were TTU NDI to determine the extent of damage. Appendix J contains the C-scans of the TTU performed on specimens 1B-3, 1B-4, and 1C-2. In these C-scans the dark square shaped region on the side of the doubler is used for setting the gain on the tester and not a delamination area. From these scans it can be seen that specimen 1B-4 (E=100 in-lbs) had a smaller delamination area than specimen 1C-2 (E=100 in-lbs) though they were both impacted with the same energy level. This is attributed to the higher bending stiffness of the 9-ply cross laminate doubler (1C-2) than the 6-ply unidirectional laminate doubler (1B-4) even though they both have similar inplane axial stiffnesses. Figures , , , are photographs of three specimens impacted at 100, 500, and 1200 in-lbs. Table relates the specimens in figures , , with the tested impact energy level. All baseline impact testing was performed at room temperature conditions. The specimens impacted at 100 and 300 inch-lbs were fatigue tested per conditions and method described in Section to help understand the durability after impact. The specimens impacted at 500 and 1200 inch-lbs were static tested as described in Section to help understand the residual strength after impact. The impact test data for the six spare baseline specimens impacted, then fatigue or static tested, is contained in Appendix K. D

94 Specimen Number Tested Impact Energy 1B inch-lbs IB inch-lbs 1B inch-lbs Table Tested Impact Energy for Specimens in Figures , , and D

95 D

96 D

97 Referring to Appendix K, the two specimens which were impacted at 500 and 1200 inch-lbs showed residual strengths of 52 and 43 ksi respectively. Typical allowable stress levels for 7075-T6 sheet are around 70 ksi. The two specimens impacted at 100 inch-lbs were fatigue tested (cracks detected outside the doubler) and failed before 300,000 cycles. The two specimens impacted at 300 inch-lbs were fatigue tested and endured the 300,000 cycles without crack detection. After fatigue testing, residual static tension strength testing was accomplished and both specimens showed the same strength, 83 ksi, which was the average of the static residual strength tests done on the specimens which endured 300,000 cycles from the fatigue testing described in Section Parametric Tests After the baseline testing was completed a number of parametric tests were accomplished, each with its own test matrix. The field repair test matrix contained in Appendix D, shows the test matrix number for each of the parametric tests. The specimens D

98 fabricated for testing were numbered according to the test matrix they would be used in. For example specimen TM10-1 was a specimen used in test matrix 10. The following sections describe the static, fatigue, and impact testing done in each test matrix in Appendix D. The parametric test data and analyses are contained in Appendix L. The drawings used to fabricate each specimen in the parametric testing are enclosed in Appendix I Additional Tests on the Baseline Geometry Six specimens were fabricated which were identical to doubler configuration 2A (geometry 2 with unidirectional plies and a normal lay-up) which was down selected from the baseline fatigue and static tests performed in Section These specimens were tested at room temperature. They are also the only specimens which were not labeled according to the test matrix where they were used, but instead are labeled TM1-1 through TM1-6. In the field repair test matrix, Appendix D, these specimens are used in test matrix 4.2. Appendix L contains the test data for these fatigue and residual strength tests. Three of the six fatigue specimens ranout without crack detection and two others were greater than 200,000 cycles when the specimen failed. One specimen failed early at 89,454 cycles, and the crack was found to start at a burr in the stop drill, see Section However, there was one specimen which detected a crack and was not failed: TM1-6. Crack growth data was recorded for this specimen and is included in Figure Change in Laminate Thickness D

99 This test matrix, 4.5 in Appendix D, was performed to evaluate the effect of two less plies than those required to match the stiffness. Three specimens were made with 4- plies instead of 6-plies on an inch thick sheet as described in the baseline testing. Appendix L contains the test data for these fatigue and residual strength tests. One of the specimens ranout, and, in two of the specimens cracks were detected before 200,000 cycles. For these two specimens crack growth data was recorded and is shown in figure No statistical difference existed between the specimens made with 4-plies versus the specimens made with 6-plies for cycles to crack detection and total cycles Crack Size in Different Aluminum Thicknesses These tests were performed to determine if differences existed between having and 0.5 inch saw cut with a stop drill and no doubler versus doublers over a 0.5 inch and 1 inch saw cut with a stop drill. Also examined, were each of these three variables on three different aluminum thickness: 0.032,0.063(baseline), and 0.10 inch. Test matrix 5 in Appendix D shows the variables used. Three specimens were replicated for each of the conditions shown. Appendix L contains the test data for these fatigue and residual strength tests. The and inch thick fatigue specimens did not crack within 300,000 cycles at the stop drill, with either the 0.5 or 1.0 inch saw cuts. However, all the 0.1 inch thick fatigue specimens showed crack detection early in the fatigue life. On the undoubled specimens, crack detection and propagation was quick. All the undoubled specimens, irrespective of aluminum thickness, experienced two piece failures before 6,000 fatigue cycles. The crack growth results for the undoubled specimens are shown in D

100 figure The crack growth information for the other doubled specimens on which cracks were detected are shown in figure The statistical analysis showed significant differences between: doubler and undoubied specimens for both crack detection cycles and total cycles; and specimens made with different thickness with a 95% probability that 0.1 inch thick will initiate cracks before inch thick specimens. For this reason a more detailed study of applying boron/epoxy doublers on 0.1 inch thick aluminum is tested in test matrix 21, see Section D

101 Environmental Exposure The baseline specimen doubler, on an inch aluminum sheet with a 0.5 inch saw cut and stop drill, was exposed to four different environments (other than room temperature) to understand the effects of different environments on the doubler system, see test matrix 7 in Appendix D. These four different environments were: -65 F/lab air, 160 F/wet, cyclic hot/cold, and solvent exposure. Three specimens were replicated for each of the conditions shown. Also, one of the environmental conditions, cyclic hot/cold, was tested on a 0.1 inch specimen fabricated like the specimens tested in test matrix 5. The environmental conditioning used in the material property tests, Section 4.1.1, is like the conditioning used in these field repair tests. The following gives a brief explanation of each environment conditioning: -65 F air conditioning was done immediately prior to and during testing. The time required to cool the specimen to -65 F was less then 5 minutes. The specimens were tested in a frame with an integral environmental chamber which used liquid nitrogen (LN 2 ) to cool it. Thermocouples were mounted on the specimens which controlled an activation unit to add LN 2 throughout the fatigue test; 160 F/wet specimens were conditioned in a chamber at 160 F and 85% RH. The moisture intake of these specimens was monitored and the moisture absorption times established from the material property specimens. Appendix B contains the moisture absorption details for these specimens. The specimens were tested, after 40 days exposure (approximately amount of time to stabilize weight gain) at 160 F; cyclic hot/cold specimens were exposed to a thermal cycle between -65 F and 160 F at a rate of 1 cyclic per hour during the 5 Hz fatigue testing. This means that if the specimens ran out, 300,000 cycles, it would be exposed to approximately 16 complete thermal cycles; and solvent exposure specimens were soaked in Skydrol hydraulic fluid for 7-days at 120 F at Boeing Material Technology Lab. After the one week soak, the specimens were plastic wrapped and delivered to Intec to be tested within 8 hours after removal from the Skydrol exposure. Referring to Appendix L, the effect of the environmental exposure on the fatigue specimens was mixed. The inch thick specimens at room temperature, with solvent D

102 exposure, and with 160 F/wet experienced runout: 300,000 cycles without crack detection. The -65 F environment exposure specimens detected cracks on four of the five fatigue tests run. This is the only exposure condition test which was replicated more than three times. Cracks were detected in two of the three specimens exposed to cyclic hot/cold. The crack growth data is shown in figure for, the inch thick specimens exposed to -65 F and the cyclic hot/cold condition. The crack growth data for the 0.1 inch thick specimens exposed to the cyclic hot/cold condition are shown in figure The statistical analysis showed a 95% probability that the specimens exposed to -65 F condition would initiate cracks before the other environments. D

103 Effects of Fasteners and Disbonds The baseline specimen configuration was modified to add fasteners and shims to different areas of the specimen. This was done to add stress concentrations from the fasteners and to simulate disbonds between the doubler and aluminum. Four different specimen configurations were used. Referring to test matrix 8.2 in Appendix D, the following is a description of each of the four configurations tested: specimen configuration B-l has two countersunk rivets installed in the aluminum, under the doubler, in line with the saw cut, see figure ; specimen configuration B-2 has three 0.5 inch diameter back-to-back Teflon shims placed between the film adhesive and the boron/epoxy doubler, see figure ; specimen configuration B-l/B-2 is a combination of fasteners like specimen B-l and shims like specimen B-2; and, specimen configuration B-3 has freeze plugs installed at the edge of the doubler corner, see figure As shown in test matrix 8.2, each specimen was fatigue tested at room temperature and replicated three times, except B-3, which was replicated twice. Also, specimen configuration B-l and B-2 were tested in cyclic hot/cold as described in Section D

104 Figure Specimen Configuration B-l Appendix L contains the test data and statistical analysis on the data. All of the fatigue specimens with rivets only, configuration B-l, room temperature and cyclic hot/cold, cracked early in the fatigue life, less than 60,000 cycles. The specimens having Teflon shim, configuration B-2, had three specimens crack and three which ranout. The crack growth information for the three specimens which cracks were detected are shown in figure The three specimens with configuration B-l /B-2 which had a combination of both rivets and shims showed the earliest crack detection, all three before 30,000 cycles. The statistical analysis in Appendix L shows that there is a 95% probability that when there are fasteners under the doubler, cracks will initiate sooner than with the other configurations tested. Crack growth information was not available on the specimen D

105 configurations with the two fasteners under the doubler and in-line with the sawcut (B-l and B-l/B-2) because they did not have enough room to install crack growth gages. Configuration B-3 had one specimen that ranout and one specimen which a crack was detected at 40,442 cycles Impact In this phase of the testing the baseline configuration specimen was evaluated to determine the effect of two different impact energies occurring at three different environments: -65 F/lab air, room temperature, and 160 F/wet, see test matrix 10 in Appendix D. The two different impact energies, 100 and 300 inch-lbs, were determined from preliminary impact testing done on extra specimens from the baseline testing, see Section These specimens were impacted identically to the specimens tested in the baseline testing section: 1 inch diameter hemispherical instrumented impactor with a 3 x 5 inch window. The specimens used in the environmental testing were conditioned like the specimens in Section All environmental conditioning was performed prior to impacting. Impact, fatigue, and static testing was done at room temperature after conditioning. All impact coupons were scanned by TTU and C-scan's, as shown in Appendix J. Six specimens were statically tested after impact, three each at 100 and 300 inch-lbs. They showed no difference versus the static results obtains in the baseline testing, see Section D

106 However, the impact specimens from Section statistically tested after impacting at 500 and 1200 inch-lbs showed a significant reduction in strength, see figure The specimens impacted at 100 and 300 inch-lbs showed no visible damage whereas the specimens impacted at 500 and 1200 inch-lbs showed visible damage. It should also be noted that the specimens exposed to the higher impact levels were of the smaller doubler geometry (T) from the baseline geometry testing. However, based on static testing done in Section , the smaller geometry (T) showed strengths equal to the other two doubler geometries, see figure Figure shows a summary of global tensile modulus for the specimens statistically tested after impact. All 18 fatigue-after-impact specimens evaluated experienced runout, irrespective of the impact energy or impact temperature (except one specimen which failed outside the doubler area). Thus no crack growth after impact information is available. This showed that impact events did not have an observed detrimental effect on the fatigue life of these specimens and maybe the opposite effect. In fact the analysis of variance (AOV) showed that there is a 98% probability that there is significant difference between crack detection for impact specimens versus non-impact specimens. When a Tukey-Kramer HSD pairwise comparison was made it showed that 100 in-lbs impact will initiate cracks later than no impact with a 95% confidence probability. The data shows that impacting the specimens prior to fatigue seems to increase the time to crack detection. However, higher impact energies (500 or 1200 in-lbs) could affect the fatigue life, based on the static tests performed. Residual strength testing was accomplished on all the specimens which ranout and are shown in Appendix L. D

107 Stop Drilling Crack Tip or Not Test matrix 13 in Appendix D examined the sensitivity to not stop drilling a crack tip. Baseline aluminum blanks (0.063 x 4 x 16 inch) with a 0.3 x inch EDM (electric discharge machine) notch were fatigued with S max of 10 ksi and R-ratio of 0.1 at 10 Hz to D

108 grow an active crack. An active crack was grown until approximately 0.50 inch long. Boron/epoxy doublers with the baseline geometry configuration were then installed on these blanks and the specimens tested in fatigue between 3-20 ksi at room temperature. Crack growth information was recorded using two crack propagation gages. The active crack specimens all propagated cracks, but the doublers stayed intact even after the aluminum totally cracked in two. It was evident that the stop drill does help in arresting or delaying crack propagation, because the baseline specimens did not always crack before 300,000 cycles. Crack growth information on these specimens is as shown in figure Residual strength static results are included in Appendix L and are significantly less, ksi, than values observed in the baseline geometry testing, ksi Changes in Plv Drop-off Rate Test matrix 14 in Appendix D examined the sensitivity to a reduction in ply drop-off rate from 1:30 used on the baseline geometry tests, to a 1:15. Since the ply thickness is approximately 5 mils, a 1:30 ply-drop gives a distance of 0.15 inches (+ 0.05) between different plies, see field repair drawings in Appendix I. Originally 1:10 ply drop-off was proposed in the phase I contract, see Section 1.2. However, after discussion with the personnel fabricating the specimens it was determined that a 1:10 ply drop-off would be difficult to fabricate and with a tolerance of inch, it could have no taper. For these reasons a 1:15 ply drop-off was used. D

109 The specimens with a 1:15 ply drop-off yield mixed results, with one experiencing runout and the other two cracking early in fatigue life, less than 100,000 cycles. This showed that the doubler ply drop-off rate could be an issue in designing an effective doubler. The AOV showed an 85% probability that the 1:15 will give different cycles to crack detection than the 1:30. Crack growth information on the two specimens which cracked, is shown in figure An investigation into the crack initiation site, see Section showed that specimen TM14-1 had a pit in the wall of the stop drill prior to doubler application which helped initiated the crack. An investigation by Textron into the crack initiation site of specimen TM14-2, showed that the crack initiated at a burr in the stop drill Changes in Cure Pressure Test matrix 17 in Appendix D examined the sensitivity to different doubler cure pressures. The cure pressure used on the baseline geometry was full vacuum, approximately 25 inches Hg. Coupon tests in Section 2.2 cured at different pressures showed that 15 inches Hg gave the best shear strength values. For this reason all the parametric test specimens were cured at a pressure of 15 inches Hg. These tests were performed to determine if differences between different cure pressures gave different results. Three cure pressures were evaluated: 5 inches Hg, 15 inches Hg, and 20 inches Hg. All the fatigue testing was done at room temperature. D

110 Eight of the nine specimens tested in test matrix 17 ranout. Only one specimen (TM17-9, 5 inches Hg) had a crack detected at approximately 200,000 cycles. The crack growth information for this specimen is presented in figure From this data, there is no statistical difference between the three cure pressures tested Restraining Lateral Bending Test matrix 18 in Appendix D examined the sensitivity to restraining the specimens to lateral out-of-plane deflection or bending. All the field repair specimens with boron/epoxy doubters on one side of the aluminum only had an inherent curvature due to the thermal expansion mismatch between the aluminum substrate and the boron/epoxy doubler (see Section ). Also, the neutral axis of the specimens is shifted through the doubler region from the center of the applied load which causes bending in the specimen during loading. Referring to the baseline testing, the S min of 3 ksi was used to reduce residual curvature and load reversal in the specimens. In test matrix 18, lateral restraint was applied at room temperature at 3 ksi. Then room temperature specimens were fatigue tested at 0-18 ksi and the -65 F/lab air exposure specimens were fatigue test at 3-20 ksi. Therefore, when the room temperature fatigue test stressed the specimen between 0 and 3 ksi, the restraint was preventing bending, and when the -65 F/lab air specimens were cooled to test temperature they were bearing on the restraint. The lateral restraint helped to prevent the thermal stress from causing the specimen to undergo a stress reversal upon cooling prior to the fatigue test. The lateral restraint used on the room temperature D

111 specimens is shown in figure and the lateral restraint used on the -65 F/lab air specimens is shown in figure D

112 Figure Photograph of Lateral Bending Restraint for Test Matrix 18, -65 F Specimens The two specimens which were laterally restrained and fatigue tested between 0-18 ksi ranout, see Appendix L. Lateral restraint significantly increased the fatigue life of the repair when this test is compared with the three specimens in test matrix 20 which were fatigue tested between 0-18 ksi without lateral restraints. The results of test matrix 20, which will be discussed in Section , were that two of the specimens initiated cracks before 75,000 cycles and one ranout. The AOV in Appendix L showed a 78% probability that specimens with lateral restraints will initiate cracks later than specimens without lateral restraints for a 0-18 ksi fatigue loading. The two specimens which were laterally restrained and fatigue tested at -65 C F between 3-20 ksi both initiated cracks before 105,000 cycles, see Appendix L. Figure D

113 shows the crack growth information for these two specimens. Lateral restraint increases the fatigue life of the repair when this test is compared with the six specimens in test matrix 7 which were fatigue tested at -65 F between 3-20 ksi, without lateral restraints. Also, the crack growth rates are less when the specimen is restrained versus unrestrained, comparing figure with figure The AOV did not show a statistically significant difference between specimens exposed to -65 F with lateral restraints and those without lateral restraints Changes to Fatigue Spectrum Test matrix 19 in Appendix D examined the sensitivity to different fatigue loading spectrums. Fatigue testing to this point has been 5 Hz constant amplitude sine wave loading. Two baseline field repair specimens were tested at two different 1 Hz frequency loading: sine wave and square wave loading, see figure Also tested in this matrix were three specimens loaded with a 767 body crown spectrum slightly modified to remove compression from the spectrum. All five specimens ranout to 300,000 cycles and gave good residual strength results between ksi. D

114 Changes in Fatigue Load Level In test matrix 20, three baseline geometry field repair specimens were fatigue tested at 0-18 ksi. The results of this test were compared with the results of laterally restrained fatigue tests in test matrix 18, see Section As previously mentioned, two of the specimens initiated cracks before 75,000 cycles and one ranout. The AOV in Appendix L shows a 87% probability that the fatigue stress of 0-18 ksi will give less cycles until crack detection than a fatigue stress of 3-20 ksi. The crack growth curves for the two specimens which initiated cracks is shown in figure D

115 inch Aluminum Investigation In test matrix 5, six specimens were tested which had a boron/epoxy doubler over a 0.1 inch thick specimen, see Section All six specimens initiated cracks early in the fatigue life, before 56,000 cycles. To further understand some of the variables which affect a doublers performance over a 0.1 inch aluminum base, test matrix 21 was created. In test matrix 5, the six specimens had an eight ply doubler on a 0.1 inch aluminum base, and were fatigue tested at 3-20 ksi. Three of these six specimen doublers were applied over a 0.5 inch long sawcut and stop drill (like the baseline geometry) and three specimen doublers were applied over a 1 inch long saw cut and stop drill, see the field repair drawings in Appendix I. In test matrix 21, the 0.1 inch thick aluminum base with the 0.5 inch sawcut and stop drill was used in all the tests. First examined was a ten ply boron/epoxy doubler fatigue tested between 3-20 ksi, to compare with the results obtained in test matrix 5. Of these three specimens tested, one ranout, and two initiated cracks before 135,000 cycles. Figure shows the crack growth information on the two specimens which initiated cracks. The AOV shows a 78% probability that 8-plys will initiate cracks sooner than 10- plys, and, the AOV for total cycles shows a 90% probability that 10-plys will have a longer fatigue life. D

116 Next examined was the same ten ply doubler just mentioned in the previous paragraph which was fatigued between 3-20 ksi and compared with the 10-ply doubler specimens fatigued between 0-18 ksi. Three of theses specimens were fatigue tested between 0-18 ksi. All three initiated cracks before 44,000 cycles. Figure shows the crack growth information on the three specimens which initiated cracks. The AOV between the two fatigue load levels (3-20 and 0-18) shows an 81% probability that 0-18 ksi stress level will initiate cracks sooner than 3-20 ksi, and there is a 91% probability that 3-20 ksi stress level will endure a longer fatigue life than 0-18 ksi, which agrees with the results in Section D

117 It is strongly believed that the influence of greater eccentricities created with a single sided doubler repairing thicker aluminum structure is the primary reason for earlier crack initiation with the 0.1 inch thick specimens. The bending stresses created by this eccentricity when the specimen is loaded are much greater than the axial in-plane stresses from loading the specimen between 3-20 or 0-18 ksi. To test this, three specimens were fabricated with 2 four ply doublers applied back-to-back, sandwiching the 0.1 inch aluminum base, see the field repair drawings in Appendix I. These three specimens were also tested between 0-18 ksi. The results from these specimens were much better than other doubler configurations on a 0.1 inch thick aluminum structure. Two of the specimens ranout and one developed a crack beyond the doubler due to a rough edge finish. These results show that bending stresses from the application of a single sided doubler have a significant effect on the performance of the repair system as the aluminum substrate's thickness increases Microscopic and Failure Analysis A preliminary investigation into the crack initiation sites and failure types of some of the specimens tested on this contract was accomplished. Table tabulates the specimens analyzed at the crack initiation site and the results of that analysis. Eight of the 22 specimens analyzed had burrs at the edge of the stop drill due to fabrication, which caused cracks to initiated. Although this information is good to know for a test program, burrs are a "fact of life" in the drilling of holes, including stop drills. This is one reason D