MIL-HDBK-3F Volume 3,Chapter 4 Building Block Approach for Composite Structures 3.The test data generated will be reduced,statistically,to obtain allowable type values using the B- basis value(90%probability,95%confidence)approach.For EMD prototypes use the guidelines in Section 4.4.1.1. 4.Develop input ply allowables for use in analytical methods that can be used in design/analysis.In general the lower of the ultimate or 1.5 x yield strength reduced values should be used for tension and compression.Edgewise shear strength ultimate values should be used for allowables when edgewise shear strength is not critical.The reduced (1.5 x yield)ultimate edgewise shear strength should be used when edgewise shear loads are critical. 5.Laminate design should be fiber-dominated by definition,i.e.,a minimum of 10%of the plies should be in each of the0°，+45°，-45°，and90 directions.For tape and fabric laminates,always input the 0 or 1-axis strength allowable values in both the 1-and 2-axis slots in the analytical methods for tensile and compressive loads.Shear inputs will be as described above.This ap- proach was first presented by Grimes in Reference 4.4.1.1.All laminates should be balanced and symmetric. A structure classification/allowables chart which defines the relationship between aircraft structure critical- ity and the allowables requirements for EMD and production is presented in Table 4.4.2.1(a).In Table 4.4.2.1(b)structural classification vs.physical defect maximum requirements are given so that the physi- cal defect size parameter varies indirectly with the aircraft structure criticality.Thus,aircraft structure criticality controls the reliability of the data (allowables)and the material quality that are necessary to support it. NOT PART OF BBA MATERIAL SCREENING SELECTION SPECIFICATION DEVELOPMENT PMC MATERIAL ALLOWABLES COUPONS Part b. BUILDING BLOCK APPROACH(BBAY TRADE STUDIES AND CONCEPTUAL DEVELOPMENT ELEMENTS-SINGLE LOAD PATH SELECTION AND PROOF OF CONCEPT AND ANALYTICAL METHODS SUBCOMPONENTS-MULTIPLE LOAD PATH STRUCTURAL VERIFICATION AND ANALYTICAL METHODS IMPROVEMENT COMPONENTS-CONTOURED,M.L.P. STRUCTURAL INTEGRITY AND FEM VALIDATION FULL-SCALE AIRCRAFT STRUCT. Figure 4.4.2.1 Aircraft structure development goals using Building Block Approach(BBA) 4-16
MIL-HDBK-3F Volume 3, Chapter 4 Building Block Approach for Composite Structures 4-16 3. The test data generated will be reduced, statistically, to obtain allowable type values using the Bbasis value (90% probability, 95% confidence) approach. For EMD prototypes use the guidelines in Section 4.4.1.1. 4. Develop input ply allowables for use in analytical methods that can be used in design/analysis. In general the lower of the ultimate or 1.5 x yield strength reduced values should be used for tension and compression. Edgewise shear strength ultimate values should be used for allowables when edgewise shear strength is not critical. The reduced (1.5 x yield) ultimate edgewise shear strength should be used when edgewise shear loads are critical. 5. Laminate design should be fiber-dominated by definition, i.e., a minimum of 10% of the plies should be in each of the 0o , +45o , -45o , and 90o directions. For tape and fabric laminates, always input the 0o or 1-axis strength allowable values in both the 1- and 2-axis slots in the analytical methods for tensile and compressive loads. Shear inputs will be as described above. This approach was first presented by Grimes in Reference 4.4.1.1. All laminates should be balanced and symmetric. A structure classification/allowables chart which defines the relationship between aircraft structure criticality and the allowables requirements for EMD and production is presented in Table 4.4.2.1(a). In Table 4.4.2.1(b) structural classification vs. physical defect maximum requirements are given so that the physical defect size parameter varies indirectly with the aircraft structure criticality. Thus, aircraft structure criticality controls the reliability of the data (allowables) and the material quality that are necessary to support it. Figure 4.4.2.1 Aircraft structure development goals using Building Block Approach (BBA)
MIL-HDBK-3F Volume 3,Chapter 4 Building Block Approach for Composite Structures TABLE 4.4.2.1(a)DOD/NASA aircraft structure classification vs.PMC allowables data requirements for EMD*and production PART A (Reference Figure 4.4.2.1) Aircraft Structure Classification Allowable Data Requirements for EMD and Production*Design Classification Description EMD*(Tape/Fabric) Production (Tape/Fabric) PRIMARY CARRIES PRIMARY AIR LOADS 1-lot materials testing Based on: 8-replicates per test type 5-lots of materials testing Fracture critical(F/C) Failure will cause loss of 8-replicates per test type vehicle ·Noncritical(N/C) Failure will not cause loss of 1-lot materials testing 4-lots of materials testing vehicle,cost critical 6-replicates per test type 6-replicates per test type replacement or repair SECONDARY CARRIES SECONDARY AIR AND 1-lot materials testing 3-lots of materials testing OTHER LOADS 4-replicates per test type plus 5-replicates per test type plus Fatigue critical(FA/C)and Failure will not cause loss of fatigue testing fatigue testing economic life critical(EL/C) vehicle,cost critical replacement or repair Noncritical(N/C) Failure will not cause loss of N/A 2-lots of materials testing vehicle 4-replicates per test type .Not cost or fatigue critical NONSTRUCTURAL NON-OR MINOR LOAD Based on 1-lot of materials testing BEARING 1.Estimates using data on 3-replicates per test type Noncritical(N/C) similar materials,or Failure/replacement minor 2.Vendor data,or inconvenience,not cost critical 3.Journals,magazines and books "For EMD,use procedure given for prototypes in Section 4.4.1 4-17
MIL-HDBK-3F Volume 3, Chapter 4 Building Block Approach for Composite Structures 4-17 TABLE 4.4.2.1(a) DOD/NASA aircraft structure classification vs. PMC allowables data requirements for EMD* and production PART A (Reference Figure 4.4.2.1) Aircraft Structure Classification Allowable Data Requirements for EMD and Production* Design Classification Description EMD* (Tape/Fabric) Production (Tape/Fabric) PRIMARY • Fracture critical (F/C) CARRIES PRIMARY AIR LOADS • Failure will cause loss of vehicle 1-lot materials testing 8-replicates per test type Based on: 5-lots of materials testing 8-replicates per test type • Noncritical (N/C) • Failure will not cause loss of vehicle, cost critical replacement or repair 1-lot materials testing 6-replicates per test type 4-lots of materials testing 6-replicates per test type SECONDARY • Fatigue critical (FA/C) and economic life critical (EL/C) CARRIES SECONDARY AIR AND OTHER LOADS • Failure will not cause loss of vehicle, cost critical replacement or repair 1-lot materials testing 4-replicates per test type plus fatigue testing 3-lots of materials testing 5-replicates per test type plus fatigue testing • Noncritical (N/C) • Failure will not cause loss of vehicle • Not cost or fatigue critical N/A 2-lots of materials testing 4-replicates per test type NONSTRUCTURAL • Noncritical (N/C) NON- OR MINOR LOAD BEARING • Failure/replacement minor inconvenience, not cost critical Based on 1. Estimates using data on similar materials, or 2. Vendor data, or 3. Journals, magazines and books 1-lot of materials testing 3-replicates per test type *For EMD, use procedure given for prototypes in Section 4.4.1
MIL-HDBK-3F Volume 3,Chapter 4 Building Block Approach for Composite Structures TABLE4.4.2.1(b) DOD/NASA aircraft structure classification vs.PMC physical defect minimum requirements for EMD*and production PARTS A AND B (Reference Figure 4.4.2.1) Aircraft Structure Physical Defect Maximum Requirements for Parts:Carbon or Glass Reinforced PMC Example Classification Description Tape Fabric PRIMARY CARRIES PRIMARY AIR LOADS s2%porosity over s5%of area.No s3%porosity over s5%of area.No delaminations allowed.No edge delaminations allowed.No edge Fracture critical(F/C) Failure will cause loss of delaminations allowed (including delaminations allowed (including vehicle holes). holes). ·Noncritical(N/C) Failure will not cause loss of vehicle,cost critical replacement or repair. SECONDARY CARRIES SECONDARY AIR ≤2%porosity over≤10%of area. ≤3%porosity over≤10%of area. OTHER LOADS No delaminations.No edge No delaminations.No edge Fatigue critical(FA/C) Failure will not cause loss of delaminations allowed (including delaminations allowed (including vehicle.cost critical holes). holes). replacement Noncritical(N/C) Failure will not cause loss of vehicle Not cost or fatigue critical NONSTRUCTURAL NON-OR MINOR LOAD BEARING s3%porosity over s10%of area. ≤4%porosity over≤15%of area. Delaminations over s2%of area. Delaminations over s2%of area. Noncritical(N/C) Failure/replacement minor Repaired edge delaminations s4% Repaired edge delaminations <4% inconvenience,not cost critical of edge length or hole of edge length or hole circumference are allowed. circumference are allowed. *For EMD,use procedure given for prototype in Section 4.4.1 4-18
MIL-HDBK-3F Volume 3, Chapter 4 Building Block Approach for Composite Structures 4-18 TABLE 4.4.2.1(b) DOD/NASA aircraft structure classification vs. PMC physical defect minimum requirements for EMD* and production PARTS A AND B (Reference Figure 4.4.2.1) Aircraft Structure Physical Defect Maximum Requirements for Parts: Carbon or Glass Reinforced PMC Example Classification Description Tape Fabric PRIMARY • Fracture critical (F/C) CARRIES PRIMARY AIR LOADS • Failure will cause loss of vehicle ≤2% porosity over ≤5% of area. No delaminations allowed. No edge delaminations allowed (including holes). ≤3% porosity over ≤5% of area. No delaminations allowed. No edge delaminations allowed (including holes). • Noncritical (N/C) • Failure will not cause loss of vehicle, cost critical replacement or repair. SECONDARY • Fatigue critical (FA/C) CARRIES SECONDARY AIR & OTHER LOADS • Failure will not cause loss of vehicle, cost critical replacement ≤2% porosity over ≤10% of area. No delaminations. No edge delaminations allowed (including holes). ≤3% porosity over ≤10% of area. No delaminations. No edge delaminations allowed (including holes). • Noncritical (N/C) • Failure will not cause loss of vehicle • Not cost or fatigue critical NONSTRUCTURAL • Noncritical (N/C) NON- OR MINOR LOAD BEARING • Failure/replacement minor inconvenience, not cost critical ≤3% porosity over ≤10% of area. Delaminations over ≤2% of area. Repaired edge delaminations ≤4% of edge length or hole circumference are allowed. ≤4% porosity over ≤15% of area. Delaminations over ≤2% of area. Repaired edge delaminations ≤4% of edge length or hole circumference are allowed. *For EMD, use procedure given for prototype in Section 4.4.1
MIL-HDBK-3F Volume 3,Chapter 4 Building Block Approach for Composite Structures 4.4.2.2 PMC composite building block structural development for DOD/NASA EMD and production air- craft Part B of the flowchart in Figure 4.4.2.1 defines the building block test effort in the general categories of: 1.Trade studies (element-single load path), 2.Selection,proof of concept,and analytical methods(sub-component-multiple load paths), 3. Structural verification and analytical methods improvement (contoured composite-multiple load path),and 4. Structural integrity and FEM validation(full-scale aircraft structure testing). The allowables shown in Figure 4.4.2.1 Part A and in Table 4.4.2.1(a)logically flow into Part B,building block testing.Table 4.4.2.1(b)on physical defect requirements applies to both Parts A and B.The Part B building block test effort is delineated in Table 4.4.2.2(a)in accordance with the part's structural classifica- tion.The four categories above are defined in detail for each structural classification.with the higher structural classification requiring more testing and analysis.The key point here is that these are guide- lines for structural development testing.The actual structural testing needed for a specific classification of structure could be more or less,depending on the vehicle's mission and whether it is manned or un- manned.Knowing the structural part classification,the aircraft's purpose and mission,risk analysis can be applied to minimize testing cost and risk.FEM and closed form composite analysis methods utilizing proper mechanical and physical properties and allowables input data will be necessary every step of the way.Failure modes and loads(stresses)as well as strain and deflection readings must be monitored and correlated with predictions to assure low risk.The use of FEM or other analysis methods alone(without testing)or with inadequate testing that does not properly interrogate failure modes,stresses(strains),and deflections for comparison with predictions can create high risk situations that should not be tolerated. Another risk issue for composite structure is quality assurance(QA),a subject that applies to both Parts A and B.Table 4.4.2.2(b)presents the nominal QA requirements for the categories of 1.Material and process selection,screening,and materials specification qualification, 2.Receiving inspection/acceptance testing. 3.In-process inspection, 4.Non-Destructive inspection(NDI), 5. Destructive testing (DT),and 6.Traceability. The QA requirements in each of these categories vary with the structural classification,with the higher classification requiring more quality assurance.By following the procedure outlined in this table,the amount of QA necessary to keep risk at an acceptable level can be ascertained.Again the amount of QA needed and the risk taken will be a function of the aircraft type and mission and whether it is manned or unmanned.Risk and cost are inversely proportional to each other for composite structural parts in each classification,so the determination of acceptable risk is necessary to this building block test program for EMD and production. 4-19
MIL-HDBK-3F Volume 3, Chapter 4 Building Block Approach for Composite Structures 4-19 4.4.2.2 PMC composite building block structural development for DOD/NASA EMD and production aircraft Part B of the flowchart in Figure 4.4.2.1 defines the building block test effort in the general categories of: 1. Trade studies (element-single load path), 2. Selection, proof of concept, and analytical methods (sub-component-multiple load paths), 3. Structural verification and analytical methods improvement (contoured composite-multiple load path), and 4. Structural integrity and FEM validation (full-scale aircraft structure testing). The allowables shown in Figure 4.4.2.1 Part A and in Table 4.4.2.1(a) logically flow into Part B, building block testing. Table 4.4.2.1(b) on physical defect requirements applies to both Parts A and B. The Part B building block test effort is delineated in Table 4.4.2.2(a) in accordance with the part’s structural classification. The four categories above are defined in detail for each structural classification, with the higher structural classification requiring more testing and analysis. The key point here is that these are guidelines for structural development testing. The actual structural testing needed for a specific classification of structure could be more or less, depending on the vehicle’s mission and whether it is manned or unmanned. Knowing the structural part classification, the aircraft’s purpose and mission, risk analysis can be applied to minimize testing cost and risk. FEM and closed form composite analysis methods utilizing proper mechanical and physical properties and allowables input data will be necessary every step of the way. Failure modes and loads (stresses) as well as strain and deflection readings must be monitored and correlated with predictions to assure low risk. The use of FEM or other analysis methods alone (without testing) or with inadequate testing that does not properly interrogate failure modes, stresses (strains), and deflections for comparison with predictions can create high risk situations that should not be tolerated. Another risk issue for composite structure is quality assurance (QA), a subject that applies to both Parts A and B. Table 4.4.2.2(b) presents the nominal QA requirements for the categories of 1. Material and process selection, screening, and materials specification qualification, 2. Receiving inspection/acceptance testing, 3. In-process inspection, 4. Non-Destructive inspection (NDI), 5. Destructive testing (DT), and 6. Traceability. The QA requirements in each of these categories vary with the structural classification, with the higher classification requiring more quality assurance. By following the procedure outlined in this table, the amount of QA necessary to keep risk at an acceptable level can be ascertained. Again the amount of QA needed and the risk taken will be a function of the aircraft type and mission and whether it is manned or unmanned. Risk and cost are inversely proportional to each other for composite structural parts in each classification, so the determination of acceptable risk is necessary to this building block test program for EMD and production