R. J. Bucci, H. Sklyut, L. Mueller and M. A. James. Alcoa, Inc. D. L. Ball. Lockheed-Martin. J. K. Donald. Fracture Technology Associates

Transcription

1 Advances in Testing and Analytical Simulation Methodologies to Support Design and Structural Integrity Assessment of Large Monolithic Parts: A New Perspective R. J. Bucci, H. Sklyut, L. Mueller and M. A. James Alcoa, Inc. D. L. Ball Lockheed-Martin J. K. Donald Fracture Technology Associates ASIP Nov 1 Dec Memphis, Tennessee Boeing 7 th Unitized Structures Conf., Nov. 10,

2 Acknowledgements: The work has many contributors over the span of 3+ decades Gary Bray John Brockenbrough Rich Brazill Pete Brouwer John Dalton Markus Heinimann Mark James Bill Kuhlman Mike Kulak Larry Mueller Mark Newborn Bob Schultz Henry Sklyut Steve Wallace John Watton Greg Wilson Alcoa Dale Ball Ralph Bush Keith Donald Lockheed-Martin USAFA (formerly Alcoa) Fracture Technology Assoc. Boeing 7 th Unitized Structures Conf., Nov. 10,

3 Abstract The quest for lighter and more affordable airframes has accelerated demand for thicker/wider/shaped alloy products (plate, extrusions, forgings, castings) and manufacturing technologies (e.g., high-speed machining, weld-joining) to grow applications of unitized structure. Capturing the full benefit of these technologies requires that residual stress effects be accounted for in both material characterization and final part design. In the case of thick or shaped metallic products, residual stresses from thermo-mechanical processing can introduce bias and large scatter effects into coupon-based durability and damage tolerance property determinations, which in turn confounds the ensuing transfer to final design. The presentation will describe efforts of Alcoa and others directed to developing improved fatigue crack growth rate data analysis methods and modeling tools that may be used to account for residual stress effects in testing and analysis. Two fundamental principles will be discussed at this seminar: advancements in fracture toughness and fatigue crack growth rate testing and analysis; and the way forward to account for the residual stress effect(s) in analysis and design of fatigue and fracture critical structures. Case study examples are presented to validate the recommended approach, and the presentation concludes with a vision for virtual design support to large monolithic part applications. Boeing 7 th Unitized Structures Conf., Nov. 10,

4 World's largest aluminum die-forgings: A380 Wing Spars (Alloy 7085) A 330 in. Boeing 7 th Unitized Structures Conf., Nov. 10,

5 Outline Overview - Residual stress effect in fracture property testing - Toughness - Fatigue crack growth Advanced test methods/residual stress correction and the associated sensitivities - Crack opening/closing methods Pre- & post-notch displacement Adjusted Compliance Ratio (ACR) method Kmax sensitivity concept - Case Studies: Alcoa 7085-T7452 Die-Forging Experience CT vs MT specimen effects Sampling & location effects Process Simulation Tools/Vision for Virtual Design Support - Coupon-to-component transfer example - Sampling effect studies Recommendations & Summary Boeing 7 th Unitized Structures Conf., Nov. 10,

6 Thick/shaped product evaluations demand caution when residual stresses are involved Complete stress relief of thick and/or shaped product forms is seldom achievable Test coupons removed from such parts are likely to also contain residual stress - Coupon isolation partially relieves the original stress, and redistributes that remaining - While the isolated coupon residual stress state is generally less severe than that of the original host, error potential in the ensuing test result can still be significant Residual stress can bias material property comparisons - Measured result may not represent the bulk material property - Property estimates can be non-conservative - Data pooling often yields overly conservative property minima (scatter effect) - Review of the literature reveals the problem is widespread The problem adversely impacts a number of promising technologies - New stress relief tempers; advanced materials - Net & near net product forms (forgings, extrusions, castings, spray form) - Low cost structure concepts (unitized and/or weld-joined structures) - Material replacement approvals for legacy aircraft and their derivatives Boeing 7 th Unitized Structures Conf., Nov. 10,

7 Fracture mechanics based damage tolerance assessment needs large monolithic parts Pedigree material input - "True" property result Free of residual stress bias, sampling and/or geometry effects Repeatable measurement - Part-to-part - Low test scatter - Coupon sampling and isolation effects understood Coupon-to-component transfer - Role of microstructure and process path understood - Host/part/coupon residual stress states understood - Crack drive solution(s) considering both internal & external forces - Process consistency controlled & understood (part-to-part) Boeing 7 th Unitized Structures Conf., Nov. 10,

8 Heat-treatable alloy residual stress profiles are strongly linked to thermal quench practice Post-quench residual stress states are generally compression at the surface and tension in the interior Center of cylinder Surface of cylinder Longitudinal CL Residual stress, ksi Radial σtan σlong σrad Tangential Distance from center of cylinder, in. Residual stress distribution of Al 7075 cylinder quenched in cold water spray Boeing 7 th Unitized Structures Conf., Nov. 10,

9 Residual stress induced clamping (or opening) can measurably impact crack tip stress intensity factor (a) Specimen location within parent slab. (b) Isolated specimen longitudinal residual stress distribution. M (c) Clamping moment developed after machining crack starter slot. M Portrayal of common crack tip clamping occurrence associated with compact specimen isolation from an unfully stress-relieved host Boeing 7 th Unitized Structures Conf., Nov. 10,

10 A simplified correction practice has been devised for fracture toughness (Kic) testing δ 1 δ 2 δ=0 P P δ = δ 2 δ 1 Measure the specimen height before and after machining the crack starter notch. Load New PQ 5% secant offset line from new origin ASTM B-909 Clamping scenario δ σ = σ δ Analyze test record with new origin displaced by δ Crack opening displacement Kic test residual stress correction schematic Boeing 7 th Unitized Structures Conf., Nov. 10,

11 Fracture toughness test results from partially stress relieved product can be misleading without proper interpretation 36 Kic, ksi(in) 1/2 (S-L, T/2) Stress relieved Non-stress relieved W Non-stress relieved, corrected Kic Specimen width, W (in.) Fracture toughness specimen size effect study (7050-T74 and -T7452 hand forging) Results showing interaction of internal stress state and specimen size on measured fracture toughness from similar 4-in. thick billets, one stress relieved and the other not Boeing 7 th Unitized Structures Conf., Nov. 10,

12 The correction practice reverses when the notch machining results in crack opening δ 1 δ 2 δ σ = +σ δ = δ 2 δ 1 Measure the specimen height before and after machining the crack starter notch. Load 5% secant offset line from new origin ASTM B-909 New PQ δ=0 P P δ Crack opening displacement Analyze test record with new origin displaced by δ Kic test residual stress correction schematic Boeing 7 th Unitized Structures Conf., Nov. 10,

13 Alcoa 7085 Forging Green Letter Re-visited Toughness data multiple shapes, multiple lots re-evaluated per ASTM B=909 CT specimens, most all excised from web locations W Location for Before and After Meas. D RK 7085-T7452 Die-Forging Toughness Interrogation 7085-T7452 Die Forging FCG Evaluation D RK C(T) Specimen H Dimension Measurements (Before and After Slot Fabrication) C(T) Specimen H Dimension Measurements (Before and After Slot Machining) H Test Lot C(T) Specimen H Dimension (in.) Notch Reference Part Number S-Number Location Orientation W: H: B: Before After (After-Before) Change MT# Engine Support A8FH L-T Opened T-L Opened Trunion A8FH L-T Closed T-L Closed Terminal Fitting A8FH Web L-T Opened T-L Opened A8FH Web L-T Closed T-L Opened A8FH Rail L-T Opened T-L Opened MT# Landing Gear Beam A8FZ17R Web L-T Closed L-S Closed A8FZ19R L-T Opened T-L Opened MT# A8FZ18R Web L-T Opened L-T Opened T-L Opened T-L Opened Boeing 7 th Unitized Structures Conf., Nov. 10,

14 Alcoa 7085 Green Letter Toughness Data Revisited 7085-T7452 Die-Forging Kq Residual Stress Correction (ASTM B-909) 12 Corrected Kq Tested Kq (ksi-in 1/2 ) Landing Gear Beam 4 Engine Support Terminal Fitting 2 Side Stay Trunion Closing -2-4 Opening δ = (δ 2 δ 1 ), in L-T test orientation Toughness correction based on C(T) specimen height change after slot introduction Boeing 7 th Unitized Structures Conf., Nov. 10,

15 Alcoa 7085 Forging Green Letter Toughness Data Revisited: Coupon size relative to host residual stress profile & failure mode can bias the result Kq (ksi in 1/2 ) Greater Correction Excised spec. intercepts greater portion of host total res. stress profile Original Kq Corrected Kq Higher Kq New failure mode, Delam. Toughening? 7085-T7452 Die-Forge Landing Gear Beam (CT Specimen, L-T orientation) Nominal Forged Section Thickness (in.) Kq Toughness w/ & w/o Res. Stress Correction per (ASTM B-909) Boeing 7 th Unitized Structures Conf., Nov. 10,

16 Residual stress influence can be greater in narrow products Machining distortion is the result of internal stress rebalance associated with metal removal A coupon's size relative to the principal dimensions of it's host determines the fraction of host residual stress profile intercepted It is thus reasonable to expect that residual stress bias will be greater from coupons having principal dimensions comparable to those of the host M M Boeing 7 th Unitized Structures Conf., Nov. 10,

17 Boeing 7 th Unitized Structures Conf., Nov. 10, 2005 FCGR correction for residual stress bias can be ascertained from the load-cod relationship Pmax (a) (b) (c) Pop = crack opening load attributed to crack closure effect (Elber). FCG occurs for P > Pop Load - P Pop Keff Pop Keff Pop Keff v P (a) Residual stress free. Pop linked to crack surface roughness. (b) Pop increase with residual stress induced clamping moment. (c) Pop decrease with residual stress induced opening moment. Crack opening displacement - v Effect of residual stress on load-cod trace

18 Precedent has been set for use of "closure-based" correction to remove residual stress bias in FCGR data (Bucci, ASTM STP 743, 1981 & Bush et. al., ASTM STP 1189, 1993) K, Mpa(m) 1/2 Keff, Mpa(m) 1/ Specimen #1 Specimen # Specimen #1 Specimen # da/dn, in/cycle da/dn, m/cycle da/dn, in/cycle da/dn, m/cycle 10-8 Uncorrected Corrected K, ksi(in) 1/2 Keff, ksi(in) 1/2 FCGR data from two partially stress relieved 7050-T7452 forgings (S-L orientation, R = 0.33, high humidity air) Boeing 7 th Unitized Structures Conf., Nov. 10,

19 Recent advances in closure correction methodology Inaccurate accounting of closure and the residual stress effect results in erroneous life prediction - Important in near-threshold region where majority of structural life is spent - Important to reconciling coupon-to-component transfer issues, particularly for thick, complex shape parts The current ASTM E647 closure measurement approach has deficits - Overly conservative, particularly Keff thresholds (high weight penalty) - Mixed agreement against small flaw and variable R-ratio data sets - High measurement variability The adjusted compliance ratio (ACR) and Kmax sensitivity methods offer significant interpretive advantages over current ASTM-E647 Kapplied and Keff based practices - Better threshold and near-threshold FCGR estimates - Better normalization of R-ratio effect - Better agreement between large and small flaw data - Better accounting of residual stress effects Boeing 7 th Unitized Structures Conf., Nov. 10,

20 Schematic of ASTM Opening Load Method P eff (ASTM) Straight-line fit to upper linear portion of the curve to determine open-crack compliance Pop defined as the point where the slope deviates 2% from the fully open compliance Assumes only that portion of the load cycle between Pop and Pmax contributes to FCG Keff = K max -K op Conceptually appealing but often problematic in practice Kop depends significantly on measurement location and distance from the crack tip and exhibits significant scatter for given location Often poor correlation between Keff and observed crack growth rates Boeing 7 th Unitized Structures Conf., Nov. 10,

21 Schematic of Adjusted Compliance Ratio (ACR) Method Developed by Keith Donald (FTA) Experimental setup, instrumentation & data collection same as ASTM E647 P eff (ASTM) Pop marks a transition point below which applied force is no longer directly proportional to crack tip strain ACR method assumes there can still be a contribution to Keff below Pop Keff is related to the actual displacement range (δcl) to the displacement range that would have occurred in the absence of crack closure (δnc) Keff = ACR Kapp CR = C s / C o C ACR = C s o C i C i Boeing 7 th Unitized Structures Conf., Nov. 10,

22 Comparison of Keff Curves Based on the ACR and ASTM Opening Load Method The ASTM opening load method typically yields lower Keff thresholds and higher near-threshold FCG rates than the ACR method. 1.0E E-02 Specimen = Middle Crack Tension W = mm; B mm R = 0.10; Frequency = 25 Hz 2ao = 5.08 mm 1.0E E-02 Specimen = Middle Crack Tension W = mm; B mm R = 0.10; Frequency = 25 Hz 2ao = 5.08 mm 1.0E T E T39 da/dn (mm/cycle) 1.0E E E-06 Keff (ACR) Keff (ASTM) Kapp Kapplied Keff - ACR Keff - ASTM da/dn (mm/cycle) 1.0E E E-06 Keff (ACR) Keff (ASTM) Kapp Kapplied Keff - ACR Keff - ASTM 1.0E K (MPa m) 1.0E K (MPa m) Boeing 7 th Unitized Structures Conf., Nov. 10,

23 FCGR Response of Alloy 2324 Plate at R = -1.0, 0.1, 0.3, 0.5, 0.7 Ref. Bray & Donald, ASTM STP 1343, 1999 Kappl basis Keff basis (ASTM) ACR method does best job of normalizing the R-ratio effect Keff basis (ACR) Boeing 7 th Unitized Structures Conf., Nov. 10,

24 K max sensitivity concept addresses the R-ratio shift associated with superposition of residual and applied forces Ref. Bray & Donald, ASTM STP 1343, 1999 Norm. K max sensitivity concept K norm = K eff (1-n) K max n At n = 0, K norm = K eff ; indicating that FCG rates depend only on K eff At n = 1, K norm = K max ; indicating that FCG rates depend only on K max K appl basis Ref. Bray & Donald, ASTM STP 1343, 1999 Figures to the right show: Plate alloy 2324 FCGR response at R = -1.0, 0.1, 0.3, 0.5, 0.7 & const. K max = 6, 9, 14 & 22 ksi in K norm basis Best normalized response uses K eff based on ACR method Boeing 7 th Unitized Structures Conf., Nov. 10,

25 ACR-Corrected Long Crack Data Matches Very Well with Small Corner Flaw Data Figure to right shows: Closure corrected long crack data (ACR method) matches the small corner flaw test result The small corner flaw data is presumed closure free in the purest sense (little crack wake to develop plasticity or roughness induced closure) da/dn (mm/cycle) da/dn (mm/cycle) 1.0E E E E E E E-07 Middle Crack Tension W = mm; B = mm R = 0.10; Frequency = 25 Hz 2ao = 5.08 mm Open Hole Corner Notched W = 38.1 mm; B= mm Hole Diameter = 6.35 mm R = 0.10; Frequency = 20 Hz Smax - gross = 82.7 MPa 7075 long & small crack data 2324 long & small crack data 7075-T7651 Keff - ACR M(T) 7075-T7651 Kapplied - Corner Crack 2324-T39 Keff - ACR M(T) 2324-T39 Kapplied - Corner Crack K (MPa m) Boeing 7 th Unitized Structures Conf., Nov. 10,

26 Representative 7085-T7452 Forging Fatigue Crack Growth Evaluation Study Objective Evaluate sample and location effects of C(T) and M(T) specimens removed from a representative 7085-T7452 die-forging Evaluate various Keff normalizing methods for removing residual stress bias from the data measurement - Closure correction (ASTM opening load & ACR methods) - Kmax sensitivity concept Boeing 7 th Unitized Structures Conf., Nov. 10,

27 7085-T7452 Die-Forging Test Program Forged Part Evaluated: Boeing 7 th Unitized Structures Conf., Nov. 10,

28 Overview of Test Program Material: T7452 Main landing gear beam die-forging - Sampling locations: web and rail Specimen type: - C(T), W = 3 in - M(T), W = 4 in Specimen orientation: L-T, T-L, S-L Stress ratio, R=0.1; Freq = 25 hz High humidity air (> 90% RH) Crack length measurement: Compliance Crack closure measurement: Compliance - ASTM opening load (2% offset) - Adjusted compliance ratio (ACR) Boeing 7 th Unitized Structures Conf., Nov. 10,

29 C(T) and M(T) Specimen Sampling Locations Landing Gear Beam Rail (805850) L-T (t/2) & L-S (t/4); CTs Landing Gear Beam Web (805852) Landing Gear Beam Web (S-No ) L-T (t/2) & L-S (t/4); MTs Landing Gear Beam Rail (805850) Sampling Locations for 7085-T7452 Die Forgings; GA Main Landing Gear Beam (X-Sect. View) Boeing 7 th Unitized Structures Conf., Nov. 10,

30 Salient findings - C(T) vs M(T) coupon FCGR response with residual stress present A unique da/dn-dk relationship is basic to fracture mechanics based life prediction (similitude concept) CT and MT specimens excised from the same host location often display notch-tip internal stress states of opposing sign (ref. prior chart) - CT notch-tip compression (clamps); MT notch-tip tension (opens), or vice-versa - FCGR inconsistencies are often missed or mistakenly labeled as a geometry effect A residual stress profile resulting in notch-tip compression (clamping) introduces an additive closure effect into ensuing crack growth measurement - Closure based K eff approaches are useful for removing residual stress bias - ACR method the preferred closure correction approach A residual stress profile producing notch-tip tension (opening) increases cyclic R-ratio and Kmax sensitivity - Useful approach to normalize Kmax sensitivity: K norm = K (1-n) eff. K n max Boeing 7 th Unitized Structures Conf., Nov. 10,

31 7085-T7452 Die-Forging FCG Evaluation C(T) Specimen H Dimension Measurements (Before & After Slot Fabrication) Lot Number S-Number Orientation Before H Dimension (in.) After (After - Before) A8FZ17R L-T (Closed) (Rail) L-S (Closed) A8FZ19R L-T (Opened) (Web) T-L (Opened) The C(T) specimen is a sensitive residual stress indicator, whereas the M(T) specimen is not. Boeing 7 th Unitized Structures Conf., Nov. 10,

32 1.0E T7452 Die Forgings: Fatigue Crack Growth R=0.1, Frequency = 25 Hz., High Humidity (RH>90%), L-T, C(T) S L-T, C(T) 1.0E-03 S L-T, C(T), DKeff (ACR) S L-T, C(T) da/dn (in/cycle) 1.0E E-05 S = Rail S = Web S L-T, C(T), DKeff (ACR) Web (L-T) CT opens Rail (L-T) CT clamps 1.0E E E-08 Large closure effect when CT clamps Closure free when CT opens DKeff (ACR) gives best agreement at low & mid K Kmax sensitivity when CT opens K K (ksi in) in) Boeing 7 th Unitized Structures Conf., Nov. 10,

33 1.0E T7452 Landing Gear Beam (Rail) Die Forgings: Fatigue Crack Growth R=0.1, Frequency = 25 Hz., High Humidity (RH>90%), L-T Orientation, 4" M(T) vs. C(T) S L-T, M(T) 1.0E-03 S L-T, M(T), DKeff (ACR) S L-T, C(T) S L-T, C(T), DKeff (ACR) 1.0E-04 da/dn (in/cycle) 1.0E-05 S = Rail Rail (L-T) MT opens Rail (L-T) CT clamps 1.0E E E-08 Large closure effect when CT clamps Closure free when MT opens DKeff (ACR) gives best agreement at low & mid K Kmax sensitivity when MT opens K -K(ksi in) in) Boeing 7 th Unitized Structures Conf., Nov. 10,

34 1.0E T7452 Landing Gear Beam (Web) Die Forging: Fatigue Crack Growth R=0.1, Frequency = 25 Hz., High Humidity (RH>90%), L-T Orientation, 4" M(T) vs. C(T) S L-T, M(T) 1.0E-03 S L-T, M(T), DKeff (ACR) S L-T, C(T) S L-T, C(T), DKeff (ACR) da/dn (in/cycle) 1.0E E E-06 S = Web Web (L-T) CT opens Web (L-T) MT clamps 1.0E E-08 Large closure effect when MT clamps Closure free when CT opens DKeff (ACR) gives best agreement Slight Kmax sensitivity when CT opens K -K(ksi in) in) Boeing 7 th Unitized Structures Conf., Nov. 10,

35 1.0E T7452 Landing Gear Beam (Rail & Web) Die Forgings: Fatigue Crack Growth R=0.1, Frequency = 25 Hz., High Humidity (RH>90%), L-T, 4" M(T) S L-T, M(T) S L-T, M(T), DKeff (ACR) 1.0E-03 S L-T, M(T) S L-T, M(T), DKeff (ACR) 1.0E-04 da/dn (in/cycle) 1.0E-05 S = Rail S = Web Rail (L-T) MT opens Web (L-T) MT clamps 1.0E E E-08 Large closure effect when MT clamps Closure free when MT opens DKeff (ACR) gives best agreement at low & mid K Kmax sensitivity when MT opens K -K(ksi in) in) Boeing 7 th Unitized Structures Conf., Nov. 10,

36 1.0E T7452 Die Forgings: Fatigue Crack Growth R=0.1, Frequency = 25 Hz., High Humidity (RH>90%), 4" M(T) vs. C(T) S L-T, C(T), DKeff (ACR) S L-T, C(T), DKeff (ACR) 1.0E-03 S L-T, M(T), DKeff (ACR) S L-T, M(T), DKeff (ACR) 1.0E-04 da/dn (in/cycle) 1.0E E E E-08 All L-T Data (ACR) DKeff (ACR) collapses data for all conditions tested - Web & Rail -CT& MT Some Kmax sensitivity at high DK (opening case) K -K(ksi in) in) Boeing 7 th Unitized Structures Conf., Nov. 10,

37 Alcoa 7085 Forging Green Letter Re-visited - Multiple shapes, multiple lots re-evaluated w/acr method - CT specimens, most all excised from web area - Findings consistent with prior study results W Location for Before and After Meas. D RK 7085-T7452 Die Forging FCG Evaluation D RK C(T) Specimen H Dimension Measurements (Before and After Slot Fabrication) H Test Lot C(T) Specimen H Dimension (in.) Notch Reference Part Number S-Number Location Orientation W: H: B: Before After (After-Before) Change MT# Engine Support A8FH L-T Opened T-L Opened Trunion A8FH L-T Closed T-L Closed Terminal Fitting A8FH Web L-T Opened T-L Opened A8FH Web L-T Closed T-L Opened A8FH Rail L-T Opened T-L Opened MT# Landing Gear Beam A8FZ17R Web L-T Closed L-S Closed A8FZ19R L-T Opened T-L Opened MT# A8FZ18R Web L-T Opened L-T Opened T-L Opened T-L Opened Boeing 7 th Unitized Structures Conf., Nov. 10,

38 da/dn (in/cycle) 1.0E E E E E T7452 Die Forging Fatigue Crack Growth Curves for Terminal Fitting, Trunion, Engine Support, and Landing Gear Beam, L-T Orientation: K Applied A8FH55, Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH56, Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Closed ( ) A8FH57; Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0001) A8FH50; Trunion L-T, Applied; H=2.4, W=2.0, B=0.25, Closed ( ) A8FH47; Engine Support L-T, Applied; H=2.4, W=2.0, B=0.25, Opened (0.0003) A8FZ18R1; Landing Gear L-T1, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0006) A8FZ18R1; Landing Gear L-T2, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0008) A8FZ17R1; Landing Gear L-T, Applied; H=3.6, W=3.0, B=0.25, Closed ( ) A8FZ19R1; Landing Gear L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0015) R = +0.1 Freq = 25Hz Lab Air L-T K-grad C(T) Specimen Highest open K Applied 1.0E E-07 Highest close 1.0E K (ksi in) Kapp, ksi in Boeing 7 th Unitized Structures Conf., Nov. 10,

39 da/dn (in/cycle) 1.0E E E E E T7452 Die Forging Fatigue Crack Growth Curves for Terminal Fitting, Trunion, Engine Support, and Landing Gear Beam, L-T Orientation: K ACR eff (ACR) A8FH55, Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH56, Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Closed ( ) A8FH57; Terminal Fitting L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0001) A8FH50; Trunion L-T, Applied; H=2.4, W=2.0, B=0.25, Closed ( ) A8FH47; Engine Support L-T, Applied; H=2.4, W=2.0, B=0.25, Opened (0.0003) A8FZ18R1; Landing Gear L-T1, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0006) A8FZ18R1; Landing Gear L-T2, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0008) A8FZ17R1; Landing Gear L-T, Applied; H=3.6, W=3.0, B=0.25, Closed ( ) A8FZ19R1; Landing Gear L-T, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0015) R = +0.1 Freq = 25Hz Lab Air L-T K-grad C(T) Specimen Highest open 1.0E E E K (ksi in) K eff ACR, ksi in KACR collapses data when notch clamped; Kmax sensitivity adjustment needed for case when notch opened Boeing 7 th Unitized Structures Conf., Nov. 10,

40 da/dn (in/cycle) 1.0E E E E E T7452 Die Forging Fatigue Crack Growth Curves for Terminal Fitting, Trunion, Engine Support, and Landing Gear Beam, T-L Orientation: K Applied A8FH55; Terminal Fitting T-L, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH56; Terminal Fitting T-L, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0001) A8FH57; Terminal Fitting T-L, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH50; Trunion T-L, Appled; H=2.4, W=2.0, B=0.25, Closed ( ) A8FH47; Engine Support T-L, Applied; H=2.4, W=2.0, B=0.25, Opened (0.0004) A8FZ18R1; Landing Gear T-L1, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0012) A8FZ18R1; Landing Gear T-L2, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0009) A8FZ19R1; Landing Gear T-L, Applied; H=3.6, W=3.0, B=0.25, Opened (0.0008) R = +0.1 Freq = 25Hz Lab Air T-L K-grad C(T) Specimen Highest open K Applied 1.0E E-07 Highest close 1.0E K (ksi in) Kapp, ksi in Boeing 7 th Unitized Structures Conf., Nov. 10,

41 da/dn (in/cycle) 1.0E E E E E E T7452 Die Forging Fatigue Crack Growth Curves for Terminal Fitting, Trunion, Engine Support, and Landing Gear Beam, T-L Orientation: K ACR eff (ACR) A8FH55; Terminal Fitting T-L, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH56; Terminal Fitting T-L, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0001) A8FH57; Terminal Fitting T-L, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0003) A8FH50; Trunion T-L, ACR; H=2.4, W=2.0, B=0.25, Closed ( ) A8FH47; Engine Support T-L, ACR; H=2.4, W=2.0, B=0.25, Opened (0.0004) A8FZ18R1; Landing Gear T-L1, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0012) A8FZ18R1; Landing Gear T-L2, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0009) A8FZ19R1; Landing Gear T-L, ACR; H=3.6, W=3.0, B=0.25, Opened (0.0008) R = +0.1 Freq = 25Hz Lab Air T-L K-grad C(T) Specimen Highest open 1.0E E K (ksi in) K eff ACR, ksi in KACR collapses data when notch clamped; Kmax sensitivity adjustment needed for case when notch opened Boeing 7 th Unitized Structures Conf., Nov. 10,

42 Crack Growth Data Reduction for Design Schematic of First Tier Validation Approach (ref. D.L. Ball, Boeing 7 th Unitized Structure Conf., 2005) Coupon Validation Boeing 7 th Unitized Structures Conf., Nov. 10,

43 Key Findings and Recommendations Residual stress can be responsible for large scatter in fracture property measurements developed from thick/large monolithic parts Guidelines and path presented to minimize the testing and interpretive problems - Warning signs and validity checks to recognize corrupted data sets - Initial proof-of-concept validated, and the technology ready for scaling to the next higher (design feature and sub-component) level - Strong advocacy is needed for upgrading test/evaluation standards Corrective protocols established to purge residual stress bias from property data - The residual stress impact on specimen crack drive can be treated as an additive (or subtractive) K-value capable of producing the equivalent residual stress induced opening (or closing) COD beyond the or residual stress free "neutral" (zero load) state. - The concept of a normalized or "master" da/dn- Keff FCGR relationship can be used to untangle the residual stress effect da/dn- Keff data with and without residual stress bias fit the same normalized FCGR relationship Best normalization results with DKeff basis the ACR method (handles crack clamping case) Further normalization improvement with Kmax sensitivity adjustment (handles crack opening case) A closure-based FCG model (e.g., FASTRAN) can be used to reverse-derive a family "true" material da/dn- K curves (free of residual stress bias) for individual R-values, which in turn can be used to predict cyclic crack growth life in the usual way Boeing 7 th Unitized Structures Conf., Nov. 10,

44 Alcoa Internal Building Block R&D Residual Stress/Machining Distortion Tool Development; Coupon-to-Component Transfer Know-how Objective: Deliver basic know-how, software, analytical methods to understand/quantify relationships between pre- and postmachining residual stresses in host product, final machined parts, test specimens and the associated impacts on material, manufacturing, and structural performance evaluations 2004/2005 Major software tool developments - Thermo-mechanical process and residual stress simulation - MDT (Machining Distortion Tool) - BARS-SP (Bi-axial Residual Stress Specimen Placement) Library of standard specimen configurations Ability to place a surface, corner or through crack into an arbitrary 3D body and/or sub-structure detail - Analytical study of residual stress effects - Tool validation against actual measurement - in process Boeing 7 th Unitized Structures Conf., Nov. 10,

45 Alcoa Vision Virtual Design Support for Monolithic Part Applications Process Simulation BLOCKY Machining Simulation MDT (Part Isolation) Coupon Isolation & Testing Simulation MDT (Coup Isolation) Arbitrary 3D Body Microstructure to Coupon Understanding Host Geometry Quenching Simulation & Stress Relief Residual Stress Prediction Microstructure Prediction Machining Distortion Residual Stress in Finished Part Microstructure in Finished Part Effect of Microstructure & Residual Stress K 1c Correction Tool da/dn Correction Tool S/N Correction Tool Coupon to Structure Understanding Residual Stress Corrected Material Properties Customer Final Geometry Geometry Residual Stress Local Microstructure Customer Support & Application Engineering Customer Loads, Requirements, and Boundary Conditions MDT (Machining Distortion Tool) Design & Certification Forged Aerostructures 3-D S/N Fatigue Small Flaw Model Location Dependent S/N Allowable 3-D Geometry Microstructure Based Strength allowable Boeing 7 th Unitized Structures Conf., Nov. 10, Crack Location Dependent Damage Tolerance Allowables

46 7xxx Aluminum Die-Forgings Old vs. New Process Old Way Residual stress & manufacturing cost improvement breakthroughs enabled via advance FEM Modeling New Way 7050-O1 Quenched Forging 7085-T7452 Cold Worked Forging (at final temper) Rough Machine to 2-4in Heat Treat/Quench/Age to -T74 Temper (No CW) Intermediate Machine (minimal flipping) Intermediate Machine (flipping, resting, etc.) Final Machine Part Residual Stresses in Double Digits (ksi) Final Machine Part Residual Stresses in Single Digits (ksi) Boeing 7 th Unitized Structures Conf., Nov. 10,

47 Alcoa MDT Tool with Automated Crack Placement Feature Redistributed σ xx As Forged Part Geometry After Machining Distortion Analysis Positioning of Crack Crack Output: - Stress - Crack Drive - Distortion Boeing 7 th Unitized Structures Conf., Nov. 10,

48 Vision for Virtual Design Support - Forging Example Host body Forging z x Residual Stress Profile in Host Body y Residual Stress - σ x (MPa) Residual Stress - σ y (MPa) Residual Stress - σ z (MPa) Boeing 7 th Unitized Structures Conf., Nov. 10,

49 Machining Part Final Geometry from the Host - Position 1 Final Mach. Part Host Forging Boeing 7 th Unitized Structures Conf., Nov. 10,

50 Machining Part Final Geometry from the Host - Position 2 Final Mach. Part Host Forging Boeing 7 th Unitized Structures Conf., Nov. 10,

51 Residual Stress in Part Final Geometry Mapped Redistributed Host Pos. 1 Host Pos. 2 Host Pos. 1 Host Pos. 2 Host Forging σ x (MPa) σ y (MPa) z x y σ z (MPa) Boeing 7 th Unitized Structures Conf., Nov. 10,

52 Distortion Vector (mm) in Part Final Geometry Host Position 1 Host Position 2 Boeing 7 th Unitized Structures Conf., Nov. 10,

53 We also want stress redistribution, distortion and K solution for cracking in the final machined part and in test coupon isolated from said part Crack Specimen Boeing 7 th Unitized Structures Conf., Nov. 10,

54 Example C(T) toughness specimen isolation problem Read in final part geometry into BARS-SP tool. Select C(T) specimen and orientation from tool bar and position in final part detail. C(T) Specimen B = 0.12", W = 1" Z Y X Boeing 7 th Unitized Structures Conf., Nov. 10,

55 Proper data interpretation requires that specimen sampling effect on residual stress induced crack drive be understood C(T) spec., W= 1" distortion vector Host Position 1 Host Position 2 Specimen Thickness Location, in Host Pos. 2 Host Pos Residual stress induced K1, ksi-in 1/2 C(T) specimen crack drive as function of final machined part location within original host Note: The relative smallness of the residual stress induced crack drive shown is attributed to smallness of the C(T) specimen (W = 1") relative to the part rib height. It is shown later that a larger W dimension would have resulted in a larger residual K. Boeing 7 th Unitized Structures Conf., Nov. 10,

56 FEA Simulation of CT Specimen Fracture Toughness Test Result shows predicted stress-strain condition in W=1" CT spec. with COD = in Host Position 1 Host Position 2 Boeing 7 th Unitized Structures Conf., Nov. 10,

57 FEA simulated load-cod diagram with & without K Q correction for residual stress; C(T) specimen (W = 1 in., B = 0.12 in.) Load (Pascal) Neutral notch position KQ 5% (w/o corr.) = 25.9 in KQ 5% (w/corr.) = 25.4 ksi in Position 2; Notch Clamped Position 1; Notch Opened KQ 5% (w/corr.) = 25.4 ksi in KQ 5% (w/o corr.) = 24.8 ksi in Both tests yield identical KQ result after correction for residual stress bias Note: the relative smallness of the residual stress correction shown is attributed to smallness of the C(T) specimen width (W = 1") relative to the part rib height. It is shown later that the needed K correction is greater for a larger W specimen. COD (mm) Boeing 7 th Unitized Structures Conf., Nov. 10,

58 Example: Sampling and location effect Specimen Size & Notch Orientation Goal: Apply BARS-SP tool to demonstrate specimen residual K dependency on postmachining stress rebalance W = 1" W = 2" W = 1" Specimen: C(T), L-T, B = 0.12", Host Pos. 1, Mid-Rib W = 2" Boeing 7 th Unitized Structures Conf., Nov. 10,

59 Specimen size & notch orientation effect on residual stress induced crack drive Notch orientation can also impact stress rebalance and the residual K as well. Specimen thickness location, in C(T) Spec., Mid-Rib L-T, B = 0.12 in Host Pos. 1 Notch clamping Notch opening W = 2" W = 1" W = 1" W = 2" Residual stress induced crack drive, K1, ksi in Boeing 7 th Unitized Structures Conf., Nov. 10,

60 Example: Sampling and location effect Placement of Final Part & Specimen within Host Machining Envelope Goal: Apply BARS-SP tool to demonstrate specimen residual K dependency on postmachining stress rebalance Host Pos. 1 Host Pos. 2 Specimen: C(T), L-T, Mid-Rib W = 1", B = 0.12" Boeing 7 th Unitized Structures Conf., Nov. 10,

61 Machined part placement within the original host also impacts the residual stress effect Specimen thickness location, in Host Pos. 1 Host Pos. 2 Host Pos. 1 Notch clamping Host Pos. 2 Notch opening C(T) spec., Mid-Rib L-T, B = 0.12 in, W = 1 in Residual stress induced crack drive, K1, ksi in Boeing 7 th Unitized Structures Conf., Nov. 10,

62 BARS-SP study to assess M(T) specimen type and effect of sample location on residual stress induced crack drive Loc.#1 Loc.#2 Loc.#4 Loc.#3 Loc.#5 Dimensions in cm Machined part excised from Pos# 1 and Pos# 2 within original host Boeing 7 th Unitized Structures Conf., Nov. 10,

63 M(T) specimen sampling study BARS-SP prediction showing variation of residual stress induced crack drive with part position (within host) and coupon location (within part) Loc.#4 Loc.#2 Loc.#1 Location #3 Web Thk., in Loc.#3 Host Pos#1 Right Crack Host Pos#2 Right Crack Host Pos#1 Left Crack Host Pos#2 Left Crack 0.05 Loc.#5 K1, ksi-in 1/ Location #1 Web Thk., in Location #4 Rail Thk., in Host Pos#1 Right Crack Host Pos#2 Right Crack Host Pos#1 Left Crack Host Pos#2 Left Crack Host Pos#1 Right Crack Host Pos#2 Right Crack Host Pos#1 Left Crack Host Pos#2 Left Crack K1, ksi-in 1/ K1, ksi-in 1/ Location #2 Web Thk., in Location #5 Web Thk., in Host Pos#1 Right Crack 0.05 Host Pos#2 Right Crack 0.05 Host Pos#1 Right Crack Host Pos#1 Left Crack Host Pos#2 Right Crack Host Pos#2 Left Crack Host Pos#1 Left Crack Host Pos#2 Left Crack K1, ksi-in 1/2 K1, ksi-in 1/2 Boeing 7 th Unitized Structures -0.6 Conf., Nov ,

64 Example: BARS-SP Tool SIF calculation for crack growth in final machined part Redistribution of residual stress (σ y ), distortion vector, and crack drive solution with crack propagation in final machined part Sensitivity to part placement in original host (Host Position# 1 & 2) Crack Deformed Mapped Boeing 7 th Unitized Structures Conf., Nov. 10,

65 Host Position #1 Redistributed Residual Stresses σ y & Crack propagation, step#1. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress K1, Psi*in^ Crack Length, in Boeing 7 th Unitized Structures Conf., Nov. 10,

66 Host Position #1 Redistributed Residual Stresses σ y & Crack propagation, step#2. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress K1, Psi*in^ Crack Length, in Boeing 7 th Unitized Structures Conf., Nov. 10,

67 Host Position #1 Redistributed Residual Stresses σ y & Crack propagation, step#3. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress K1, Psi*in^ Crack Length, in Boeing 7 th Unitized Structures Conf., Nov. 10,

68 Host Position #1 Redistributed Residual Stresses σ y & Crack propagation, step#4. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress K1, Psi*in^ Crack Length, in Boeing 7 th Unitized Structures Conf., Nov. 10,

69 Host Position #1 Redistributed Residual Stresses σ y & Crack propagation, step#5. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress K1, Psi*in^ Crack Length, in Boeing 7 th Unitized Structures Conf., Nov. 10,

70 Host Position #2 Redistributed Residual Stresses σ y & Crack propagation, step#1. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress Crack Length, in K1, Psi*in^0.5 Boeing 7 th Unitized Structures Conf., Nov. 10,

71 Host Position #2 Redistributed Residual Stresses σ y & Crack propagation, step#2. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress Crack Length, in K1, Psi*in^0.5 Boeing 7 th Unitized Structures Conf., Nov. 10,

72 Host Position #2 Redistributed Residual Stresses σ y & Crack propagation, step#3. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress Crack Length, in K1, Psi*in^0.5 Boeing 7 th Unitized Structures Conf., Nov. 10,

73 Host Position #2 Redistributed Residual Stresses σ y & Crack propagation, step#4. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress Crack Length, in K1, Psi*in^0.5 Boeing 7 th Unitized Structures Conf., Nov. 10,

74 Host Position #2 Redistributed Residual Stresses σ y & Crack propagation, step#5. Distortion vector (deformation scale = 40). K1(max value) diagram due to Residual Stress Crack Length, in K1, Psi*in^0.5 Boeing 7 th Unitized Structures Conf., Nov. 10,

75 Summary: Max. value of residual stress induced K1 vs. crack length Crack Deformed Mapped K1, ksi-in 1/2 0 Host Pos Crack Length, in Host Pos Boeing 7 th Unitized Structures Conf., Nov. 10,

76 Summary Accounting for residual stress effects in damage tolerance evaluations will be essential to thick monolithic part design and approval - Failure to account for closure and residual stresses accurately will result in erroneous conclusions and enormous inefficiencies in design, analysis and certification processes - Even stress relieved product forms can be greatly affected Path established to minimize testing & data interpretation problems - Warning signs and validity checks available to identify potential for data corruption - Test/analysis methodology established to purge residual stress bias from property data Experimental protocol established and automated Kic and da/dn data reduction and analysis practices established Correction practice has basis in crack opening/closing methodologies - A significant body of data has been generated to support initial proof-of-concept Coupon-to-component transfer path emerging - The technology is ready for scaling to the design feature and sub-component levels - Advance process simulation tools offer valuable new insights for characterization, analysis and design of large monolithic parts - Strong advocacy is needed for upgrading industry test/evaluation standards & data bases Boeing 7 th Unitized Structures Conf., Nov. 10,

77 The Aeroelastic Design and Testing of the F/A-22 by William D. Anderson Lockheed Martin Aeronautics Company Marietta, Georgia Presented at the 2005 USAF Aircraft Structural Integrity Program Conference Memphis, Tennessee 29 November - 1 December by Lockheed Martin Corporation. Lockheed Martin Aeronautics Company

78 Presentation Topics Flutter Development, Analysis Scope and Procedures Flutter and Aeroelastic Design Criteria Analysis Methods Aeroelastic Tailoring Analysis Approach Flutter Issues with Early F/A-22 Design Initial Modes of Concern Initial Trades / Aeroelastic Design Optimization Results Design for Transonic Buzz and LCO Flutter Model W/T Results / Issues Aeroelastic Design Impacts Verification, Validation and Certification Lab, Ground, and Flight Testing Certification Analysis and Documentation 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 2

79 Overview of the F/A-22 Flutter Development Aero Data Analysis Doublet Lattice Zona 51 L3 Press. Model Flutter Model Team FEM Stiffness Matrix Delta Ks Delta Ms Team Mass Data Actuator Data Mass Matrix Stiffness Response Actuator Bench Test Stores Data Geometry Mass Stiffness Aeroelastic Tailoring Flutter Analysis ASE Analysis YF-22 Design and Ground & Flight Testing Requirements to Team: Stiffness Actuator & Structural Freeplay Control Law Filters Geometry Etc. Criteria (Circa 1991) 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 3

80 YF-22/PAV Design, Ground and Flight Testing Influences YF-22 design basis for initial control loop stiffness sizing From YF-22 Testing Identified 11 Corrections or Improvements for the Flutter Excitation System. Lessons Learned applied to EMD Flutter Excitation System Design. Significant Horizontal Tail Journal Bearing Friction Effects Identified on YF-22 GVT. Result led to incorporation of low friction Horizontal Bearings on EMD Flight Flutter Test Aircraft. Only Subsonic Flutter Test Data Obtained due to Schedule Constraints. Lack of YF-22 Supersonic Data put emphasis on need for a Supersonic Flutter Model for EMD. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 4

81 Aeroelastic Design Criteria Classical Flutter The Air Vehicle, including for all probable failures, shall be free from flutter or other Aeroelastic Instabilities to 1.15 V L at constant altitude and at constant Mach. Probable failures included any single hydraulic system failure. Damping The minimum damping of any potentially critical flutter mode shall be greater than the lesser of 0.03 or 1 percent above the GVT measured mode damping. Transonic Buzz and LCO Control surfaces shall be free from buzz or LCO. A tailored criteria was developed and applied to the F/A-22 design. Aeroservoelasticity Any potential aeroservoelastically critical mode shall have a gain margin of 6dB, and separately a phase margin of +/-60 degrees. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 5

82 Structural & Aeroelastic Design Vehicle / Airframe Design Team Airframe IPTs Design Layouts Structural Sizing Design Drawings Detail Design Schedule Inputs from all Disciplines Manufacturing Maintainability Weights Aero/Thermo Structures Etc. Sub-optimization Airframe A & I Team Structures A & I Team Airframe Design Integration / Coordination Design Scheduling 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company Structural Criteria, Policy & Methods Air Vehicle Loads Team Finite Element Model Air Vehicle Aeroelastic Analysis Aeroelastic Optimization Stiffness & Freeplay Requirements Filters for ASE Materials & Processes Structural Development Tests Internal Loads Allowables Vibration & Acoustics Requirements & Sonic Fatigue 6

83 Air Vehicle Air Vehicle Finite Finite Element Element Model Model 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 7

84 Flutter Analysis Scope Gross Design Sensitivities Trend Data / Parametrics Detail Design Sensitivities Major Strength & Stiffness Updates Actuators Rudder Plies Skin Offsets Fin Spars Horizontal Shaft Hinge Placement Super Element Factors L.E. Flap Engine Boom EI & GJ Component Substitution Control Loop Stiffness Rudder Plies Fin Ply Sweep Angle Horizontal Ply Sweep Angle Horizontal Ballast Weight Boom Super Element Access Panel Effectiveness Wing Ply Sweep Internal Fuel Pylon Stiffness Store Loadings H. S. Skins Rudder Skins Flaperon Skins Wing Skins Fin & Rudder Spars & Ribs Boom Panels & Bar Elements Control Loop Stiffness Freeplay! Wts at Asets Missile Adapter Stiffness, Damping & Geometry 8 Major A/V FEM Updates 9 Less Major A/V FEM Updates Numerous A/V Aeroelastic Sizing Updates 5+ Engine FEM Updates 6+ Pylon and Missile Adapter FEM Updates 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 8

85 F/A-22 Aeroelastic Tailoring / MDO Analysis Process Define Constraints and Objectives Flutter Speeds Hump Mode Damping Ply Stack-up Strength Etc. Define Design Variables (800+ Design Variables Used) Compute Sensitivities/Derivatives Perform Optimization to Achieve Objectives for Minimum Weight Design with Known Constraints. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 9

86 F/A-22 Aeroelastic Tailoring / MDO Analysis Process Key Elements of the Aeroelastic Tailoring Process: Parametric Analysis to Develop Understanding of Issues. Tools to make Rapid Selection and Definition of Design Variables. Meaningful Constraint Definitions Tightly Coupled to Design. Ability to Compute Accurate Design Sensitivities. Ability to Rapidly Update the Air Vehicle FEM and Mass Data both for Sensitivity Analysis and Aeroelastic Resize Analysis. Optimizer Capable of Handling many Design Variables, Sensitivities, and Constraints. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 10

87 Aeroelastic Design Issues and Flutter Critical Modes Classical Flutter Wing, Fin-Rudder, Horizontal and Empennage Hump Modes Coupled Horizontal and Flaperon Transonic Buzz All Control Surfaces except Horizontal; Rudder most Critical LCO Transonic and edge of envelope (all surfaces including horizontal) Drivers - control surface inertia & stiffness and control loop stiffness and freeplay. Aeroservoelasticity Large, high inertia control surfaces coupling with vertical and roll / lateral modes of the fuselage. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 11

88 Flutter Critical Modes Early EMD Configuration Flutter Boundaries Potential for Transonic Buzz / LCO (All surfaces except Horizontal - Rudder most critical) V L 1.15 V L 52 Hz High Frequency Fin / Rudder Tip Mode Region of Potential Hump Modes Altitude 30+ Hz Rudder Rotation Modes 0 Wing Bending / Torsion Mode Mach Increased Flutter and LCO Criticality 30+ Hz Horizontal Rotation Modes 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 12

89 Initial Trades & Design Optimization Results - 52 Hz Coupled Fin / Rudder Tip Mode Flutter Boundaries V L 1.15 V L Altitude Design Trades & Tailoring: 0.25 increase in t/c Stabilized Mode but did not eliminate mechanism. Adverse Aero Impact Upper hinge bearing lowered 6 inches Eliminated flutter mechanism. 35 % Tip chord reduction Favorable for all fin-rudder modes & loop stiffness. Adverse LO Impact. Aeroelastic Tailoring - Difficult to improve mode with tailoring alone 0 Early Strength Design 0.25 increase in t/c Mach 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 13

90 Initial Trades / Design Optimization Results - 30 Hz Coupled Fin / Rudder Mode Flutter Boundaries V L 1.15 V L Altitude Initial Strength Design with Initial Control Loop Stiffnesses 0 Aeroelastic Tailored Design with Reduced Control Loop Stiffnesses Mach Aeroelastic Tailoring very effective 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 14

91 Initial Trades & Design Optimization Results Hz Coupled Horizontal Rotation Mode Primary stability drivers Control Loop Stiffness Control Surface Moment of Inertia (MOI) Tail Boom torsion-plunge coupling (Actuator access door) Skin stiffness distribution Impact on Design Design changes to minimize MOI. Control Loop Stiffness improvements A change to the actuator access door location Effectiveness of actuator access door due to maintainability loose (high clearance) fit fastener requirements was an issue. With door on bottom of tail boom, tail boom pitch-plunge coupling destabilized this mode. Parametric trades showed moving the door to the inboard side of tail boom eliminated the adverse coupling. Design change implemented. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 15

92 Design for Transonic Buzz and LCO Empirically Based Reduced Frequency, ωc/v M = 1.2, V = 220 KEAS, Vtrue = 1163 ft/sec LEGEND: Aileron Flaperon Rudder Rudder Most Critical Buzz Empirical Based Design Requirement max Freeplay max Freeplay max Freeplay ωc/v > % Loop Stiffness K effective /K nominal 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 16

93 Flutter Model Vertical Tail Assembly Installed in 4X4 Tunnel Mach Dynamic Pressure Magnitude Accelerometer Response Flutter Vertical Fin in Tunnel 1000 Mach 1.41 Q=2710 psf at Flutter Freq = 240 Hz Single Degree of Freedom Flutter Time ~ Seconds Rudder Tip Failure Total Fin Failure 500 Rudder Accel ~gs Increasing Q Failures after Flutter Time ~ Seconds 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 17

94 Flutter Model Key Result / Issue Flutter Model Key result / issue Transonic Buzz of rudder measured at an ωc/v= 0.66 Buzz Criteria used for design - ωc/v> 0.40 To redesign to ωc/v> 0.66 would have been a major impact. Approach Taken Used ENSAERO with 10 modes to correlate to flutter model at model scale. ENSAERO then ran at full scale resulting in ωc/v= 0.47 at buzz. Result indicated significant Reynolds number / scale effect. With this result, decision was made to proceed with the then current design into flight test. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 18

95 Multiple Team FEM and Aeroelastic Analysis Requirements Updates FEM Models Block I Model 638 Model 639 Model 640 Model 641 Model 642 Model 644 Model A645 Block II Model A645A Requirements Updates Control Loop Skin Sizing Backup Structure Freeplay Test Data / Correlation L-3 W/T Pressure Model Flutter Model Actuator Bench Test updates 1 update 3 updates 2 updates 3 updates Multiple Aeroelastic Sizing Updates Performed with each FEM update 3 updates Steady Aero Correlation used to update Unsteady Flutter Correlation Failure Modes Stiffness 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 19

96 MDO F/A-22 Structural Design Flutter Design Impacts Overview ~ 245 Lbs for Flutter Design Impacted by Flutter Control Loop Stiffness & Freeplay Requirements (Substructure & Actuator Design) Skin & Spar Sizing for Bending And/or Torsion Stiffness At Ply Level for Composites 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 20

97 Summary Aeroelastic Design The F/A-22 Presented many Aeroelastic Design Challenges significant impact of LO on the control surfaces and edges. Several critical modes identified and were successfully addressed Required an Integrated Approach using Parametric Analysis and Aeroelastic Design Optimization Parametrics key in identifying Hinge and Access Panel location changes. Optimization used extensively for ply and substructure changes. Aeroelastic design optimization was successfully applied to achieve a minimum weight aeroelastic design for the F/A-22. A rigorous process in place of Mil-Std Criteria was used to establish freeplay allowables for the F/A-22 for control of LCO and Buzz. The integrated structural dynamic unsteady CFD analysis (ENSAERO) showed significant Reynolds Number effects for transonic buzz, with the wind tunnel scale result being very conservative. Multiple filters developed and incorporated to address aeroservoelastic stability. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 21

98 Flutter Development / Certification Overview Aero Data Design (Circa 1991) Analysis Doublet Lattice Zona 51 L3 Press. Model Flutter Model YF-22 Design and Ground & Flight Testing Team FEM Aeroelastic Tailoring Stiffness Matrix Delta Ks Requirements to Team: Airframe Stiffness Actuator Stiffness Free play Control Law Filters Geometry Etc. Team Mass Data Mass Matrix Delta Ms Flutter Analysis Actuator Data Stiffness Response Actuator Bench Test ASE Analysis Stores Data Geometry Mass Stiffness Criteria / Spec V&V & Final Certification Verification Testing: SICs MOI Stiffness & Free play GVTs Flight Flutter Structural Coupling Final Certification: Correlation & Analysis Updates Final Analysis Certification Documentation Final Reports 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 22

99 Flutter Verification / Certification Process Preliminary Block II Flutter and ASE Analysis and Reports Basis of TISs, ETRs, AOLs, and Inspections Actuator Impedance (ETR HB9205) Aircraft SIC (TIS ST0960) Horizontal Stab. SIC (TIS ST0960) Control Loop Stiff & F.P. (TIS ST0950 Control Surf Inertia (ETRs) Store MOI & Cant Pylon GVT (ETRs) Update Vibration Predictions Air Vehicle GVT (TIS ST0930) Structural Coupling (TIS FQ0900) FES Ground Test (TIS ST0940) Vibration Correlation & Update Flutter & ASE Analysis Update Flutter, Buzz & LCO Predictions Flight Flutter Tests (TISs ST0010 & ST0080) Interim Limitations and Letter Reports 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company Final Limitations, Certification Reports and FSMP and IAT Updates 23

100 Ground Tests Actuator Bench Impedance Tests Aileron, rudder, and horizontal tail actuators tested Flaperon not tested due to similarity to rudder & aileron actuators Powered and failed (Hydraulic System) conditions tested Variations in hydraulic fluid temperature, mean load, oscillatory load, and stroke position were tested Results: The actuator stiffness less than predicted in the failed / compensator mode for several of the test conditions. Resulted in following flight manual requirements: To slow to speed << than VL after any single failure condition. To slow to final approach speed when the compensator depleted ICAW enunciates. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 24

101 Ground Tests (Continued) Structural Influence Coefficient Test Aircraft 4001 and 4003 tested Included both distributed and for point loads Point Loads on the wings, leading edge flaps, ailerons, flaperons, vertical tails, rudders, and horizontal tails. Load and deflection measurements were analyzed to obtain the Structural Influence Coefficients (SIC). A separate SIC test was conducted on isolated cantilevered horizontal tails 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 25

102 Ground Tests (Continued) Control Surface Stiffness & Freeplay Tests Tests conducted on several aircraft to verify control loop freeplay and stiffness values. Tests were conducted with nominal actuators and with solid rods of known stiffness so that the backup structure stiffness could be determined. Freeplay for nominal control loop and control loop set to maximum freeplay was measured. Tests used to calibrate field freeplay check procedure. Procedure has been incorporated into the Dash 6 / Tech Orders. 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 26

103 Ground Tests (Continued) Full Aircraft Ground Vibration Tests (GVTS) Tests were conducted on Aircraft 4001, 4003, 4005, and 4008 clean wing and with external stores. Tests used to measure / verify the total aircraft structural vibration mode shapes, frequencies, and damping values. Individual control surfaces, doors, stores and launchers, and landing gear were also tested. Ground Vibration Test Setup 29 Nov 01 Dec, 2005 USAF ASIP Conference Lockheed Martin Aeronautics Company 27