Non-Export Controlled Information Releasable to Foreign Persons Aircraft Level Dynamic Model Validation for the STOVL F-35 Lightning II David A. Boyce Flutter Technical Lead F-35 Structures Technologies Lockheed Martin Aeronautics Company Robert J. Burt Director F-35 Structures Development Lockheed Martin Aeronautics Company 2009 Aircraft Structural Integrity Program (ASIP) Conference December 1-3, 2009 Jacksonville, FL DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. 2009 Lockheed Martin Corporation 1
Tri-Variant Joint Strike Fighter (JSF) 3 Airplanes : 1 Development & Certification Team Convention Take- Off and Landing Short Take- Off and Vertical Landing Carrier Variant F-35A CTOL Span........................ 35 ft / 10.67 m Length..................... 51.4 ft / 15.67 m Wing area.................. 460 ft 2 / 42.7 m 2 Combat radius (internal fuel) >590 n.mi / 1,093 km Range (internal fuel)..... ~1,200 n.mi / 2,222 km F-35B STOVL Span....................... 35 ft / 10.67 m Length.................... 51.2 ft / 15.61 m Wing area.................. 460 ft 2 / 42.7 m 2 Combat radius (internal fuel) >450 n.mi / 833 km Range (internal fuel)...... ~900 n.mi / 1,667 km F-35C CV Span........................ 43 ft / 13.11 m Length..................... 51.4 ft / 15.67 m Wing area.................. 668 ft 2 / 62.06 m 2 Combat radius (internal fuel) >600 n.mi / 1,111 km Range (internal fuel)..... >1,200 n.mi / 2,222 km 2
F-35 Structural Certification Plan for Flutter Pre-Test Analyses As-Designed A/C FEM Actuator Bench Tests Control Surface Mass Properties Tests Follow-On / Post-1 st Flight Ground Tests Validate Additional Capability Component Ground Tests Validate Control Surface Models Aircraft Ground Tests Validate Air Vehicle Model Control Surface Stiffness & GVT Tests Control Surface Freeplay & Rigidity Tests Aircraft Modal Data from Structural Coupling Test and/or Ground Vibration Test FEM Correlated to Full Range of Design Configurations Flight Test Verification F-35B (STOVL) Complete To This Point Initial Flight Clearance Airworthiness Envelope Final Flight Test Verified Aeroelastic Stability Analysis Confirms All Margins are Met or Exceeded 1 st Flight Final Aeroelastic Stability Certification 3
High Fidelity Air Vehicle FEM Air Vehicle Finite Element Model (FEM) Employs High Fidelity Representation of Structure Element size is typically <3 inches All major skins, bulkheads, spars, ribs, access panels and doors are realistically modeled Many caps, stiffeners, and lugs are represented with SHELL elements Fewer ROD, BAR, BEAM elements are used F-35B STOVL Air Vehicle FEM >250,000 elements >180,000 grids >1,000,000 DOF Advantages: Higher resolution stiffness load paths and mass distribution Easily models representative structural change impacts to vehicle dynamics Excellent correlation to ground vibration test measured modes Example of High Fidelity AV FEM Modeling Left Horizontal Tail Substructure - Larger caps modeled with SHELLS - Hinge lug details represented - Local stiffeners represented on perimeter spars & ribs Challenges: FEM databases are huge (>1 GB) Preparing model mass distribution is complex and requires meticulous bookkeeping Lots of bookkeeping to maintain consistent node numbering with aircraft and numerous pylons, racks and stores 4
Air Vehicle FEM Validation Correlation Of Dynamics FEM To Actual Aircraft Accomplished In Build-Up Fashion Each Step Was Used To Update And/Or Verify The Accuracy Of One Or More FEM Components Component Tests Flight Control Actuators Dynamic Stiffness Testing Control Surfaces Mass Properties Measurements Static Stiffness/Deflection Measurements Modal Survey (GVT In Fixture) Aircraft Tests Freeplay & Rigidity Measurements Control Surface Actuation Loop Stiffness And Joint Slop (Freeplay) Structural Coupling Tests (SCT) Flight Control Sensor Responses Due To Control Surface Motions Ground Vibration Tests (GVT) Modal Survey Of Aircraft In Multiple Configurations 5
Actuator Dynamic Stiffness Test F-35 Actuator Impedance Test Data Requirement Unbalanced system, compression, 80 deg 80º F fluid Unbalanced system, tension, 80 deg Balanced system, compression, 80 deg Balanced system, tension, 80 deg Unbalanced system, compression, 250 deg Unbalanced system, tension 250 deg Balanced system, compression, 250 deg Balanced system, tension 250 deg Stiffness (lb/in). Stiffness Requirement 250º F fluid Stiffness Requirement box (stiffness vs. frequency) determined at very early stage in F-35 development with empirical and analytical flutter trade studies. Stiffness requirement is part of Performance Based Specification for actuator vendor. Exceeding the requirement (more stiff) is a desirable trait. Room temp (80º) data used for ground test predictions. Hot fluid (250º) data used for aeroelastic stability predictions. Frequency (Hz) 6
Control Surface Mass Properties Measurements In-House Test Equipment Obtain Mass Properties Just Prior To Installation On Aircraft Measurements Taken Weight Center-of-Gravity Mass Moment-of-Inertia About Hingeline Scope All Control Surfaces For: Every Development Flight Test Aircraft First 16 Production Aircraft of Each Variant Total: >450 Control Surfaces Horizontal Tail 7
STOVL Rudder Component GVT Rudder FEM was within 1% of measured value Stiffness tuning of middle hinge pin used to match 1 st bending mode No other FEM changes needed Mode Frequency Test FEM % difference of FEM to Test mode frequency -3.5% 2.4% -0.9% MAC* 0.99 0.94 0.98 Rotation 1st Bending 2nd Bending * Modal Assurance Criterion 8
STOVL VT & Rudder Component GVT Load. Deflection Dummy actuator load-deflection test data exhibits non-linear stiffness at low load due to threaded rod-ends Rudder Rotation mode tested at higher frequency than FEM Attributed to non-linear stiffness of threaded rod-ends on dummy actuator Deferred any tuning pending aircraft GVT with real actuator In the end, no FEM change was required Mode Frequency MAC* Test FEM % difference of FEM to Test mode frequency -3.2% * Modal Assurance Criterion -8.7% 0.97 0.94 0.96-1.0% -0.3% 0.93 VT 1st Bending Rud Rotation VT 2nd Bending VT 1st Torsion 9
Effect of Hingeline Aerodynamic Seals Plate & spring elements added to simulate seals along hingeline Sizing tuned manually and with NASTRAN SOL 200 Optimization Hingeline seals introduce significant friction and cause frequencies to rise Seals modeled as plates (CQUAD) cantilevered at VT trailing edge spar and connected by springs (CELAS) to Rudder leading edge spar Element properties tuned via manual adjustments and NASTRAN SOL 200 Optimization to improve GVT correlation Mode Frequency Test - No Seals Test - w/seals FEM - w/seals Frequency rise due to seals 1.6% 2.2% % difference of FEM to Test mode frequency 2.6% 9.8% VT 1st Bending Rud Rotation VT 2nd Bending VT 1st Torsion 10
Freeplay & Rigidity Tests Rigidity = slope [in-lbs/rad] ( Loop Stiffness ) Applied hinge moment via pneumatic actuator and fulcrum mechanism Freeplay = gap [deg] between positive & negative moment Rotation measured by differential deflections along chordwise station Live actuator required to get full system stiffness and hold surface in position Actuator position feedback logic caused dithering of deflection measurements Dithering removed by performing test over long period (>4 minutes) and applying a low-pass filter to smooth data 11
Aircraft Structural Coupling Test (SCT) These Plots Display The Transfer Function Of Sensor Response Due To Surface Command Describes Propensity Of Structural Response To Be Fed Back Through Control System Too Large Of Response Could Create An Unstable Feedback Loop Resulting In Large Dynamic Loads Magnitude (db) 60 40 20 0-20 -40-60 -80 HT-to-Nz Analytical Measured OL_PA_35 5 10 15 20 25 30 35 40 Frequency (Hz) SCT Data Exhibits Good Dynamic Response Matching And Low Feedback Gains Analytical Data Obtained From Ground Test Verified Air Vehicle FEM F-35B Structural Coupling Stability Margins Exceed Requirements Magnitude (db) 60 40 20 0-20 -40-60 -80 Flap-to-Pb (Otbd) Analytical Measured OL_UA_27 5 10 15 20 25 30 35 40 Frequency (Hz) 12
Aircraft Ground Vibration Test (GVT) STOVL GVT & Flight Clearance Completed In 2 Stages Similar Approaches Being Used For CTOL And CV Variants First STOVL Aircraft Before Variant 1 st Flight One Configuration: No Simulated Actuator Failures, No Fuel, No Stores Demonstrated No Gross Errors In Structural FEM High Fidelity Dynamics FEM Used To Show Robust Flutter Behavior For Failure States And Fuel Effects To Reinforce Confidence Sufficient For Limited Flight Envelope To Complete 1 st Flight STOVL Flutter Test Aircraft After Variant 1 st Flight Multiple Configurations Multiple Simulated Actuator Failure States Various Internal Fuel States Multiple Internal Weapon Configurations Provided High Quality Correlation Evidence Across Full Range Of FEM Configurations Sufficient To Provide Clearance To Begin Envelope Expansion Testing To Design Limits 13
Aircraft Ground Vibration Test (GVT) F-35B STOVL Flutter Test A/C Clean Wing GVT Conducted on Flight Line at LM Aero Ft. Worth Underside Soft Support System Installed at Jack Points Permits Testing with Gear-Up (simulates Up & Away, free-free conditions) Rigid Body Modes <25% of Lowest Flexible Mode Frequency 14
GVT Instrumentation & Shakers 300+ Accelerometers Recorded Simultaneously - Airframe - Engine - Landing Gear - Internal Stores - Soft Support System GVT accels (not glued to surface yet) V-Tail Shaker Wingtip Shaker H-Tail Shaker 6 Shakers Simultaneously (Random Input) - Left & Right Wing Tips - Left & Right Vertical Tails - Left & Right Horizontal Tails 2 Shakers Simultaneously (Sine Input) - Both Flaps - Both Rudders - Both HTs 15
Clean Wing GVT FEM Correlation Excellent Test-Analysis Correlation Observed Primary Mode Frequencies Agree Within <4% Modal Assurance Criteria (MAC) ~0.9 For Mode Shape Of Primary Surface/Component Moving Test FEM -2.9% Mode Frequency. % difference of FEM to Test mode frequency 3.1% 0.5% -1.4% 0.0% 2.3% 2.6% 3.1% -1.1% MAC* 0.88 1.00 1.00 0.96 0.92 0.94 0.83 0.87 0.96 Fuselage Vertical Bending Wing 1st Bending Vertical Tail 1st Bending Horiz Tail Rotation Flaperon Rotation Wing 2nd Bending Horiz Tail 1st Bending Rudder Rotation Vertical Tail 1st Torsion Test Results Verify The Dynamics FEM Accurately Represents F-35B Air Vehicle 16
External Stores Component GVT Air-to-Air Pylon, Launcher, Missile Fixture Mounted GVT of Air-to-Air Missile Dynamics FEM of Pylon, Launcher, and (flexible) Missile Test revealed non-linear behavior in 2 of 3 pylon-to-wing attachments Test FEM % difference of FEM to Test mode frequency 0.2% Caused by friction along the attachment interface FEM correlated to test when attachments were forced to react loads in directions previously modeled as free Mode Frequency MAC* 0.7% 2.0% 0.97 0.93 0.89 Lateral Yaw Pitch * Modal Assurance Criterion 17
External Stores Component GVT Air-to-Ground Pylon and Store Fixture Mounted GVT of Pylon and Store Dynamics FEM of Pylon and Store Component GVT conducted with a range of stores from: Low Weight Low Inertia Class to High Weight High Inertia Class Final pylon FEM updating being done in conjunction with aircraft external stores GVT correlation Mode Frequency Test FEM 4.3% Heavy Store Configuration (Preliminary Data) -1.4% % difference of FEM to Test mode frequency 0.6% Lateral Yaw Pitch 18
External Stores Aircraft GVT External Stores Dynamics FEM Test FEM Heaviest Store Configuration (Preliminary Data) % difference of FEM to Test mode frequency 2.3% F-35B External Stores GVT Conducted on Flight Line at LM Aero Ft. Worth Mode Frequency. 3.4% -3.3% -2.4% -9.2% 2.2% 4.2% MAC* 0.74 0.48 0.27 0.10 0.73 0.53 0.71 Air-to-Air Lateral Inb d Store Lateral Outb d Store Pitch Inb d Store Pitch Wing Torsion Outb d Store Lateral Wing Bending Preliminary Test Results Indicate Excellent Correlation For Dynamics Stores FEM 19
F-35B Envelope Expansion Cleared Envelope 1st Flight and Initial Airworthiness Testing ~50% M L Restricted Envelope Operations Beyond Initial Envelope not Authorized Design M L Simple GVT Prior To Variant 1 st Flight - Supports Limited Initial Flight Envelope - Maintains Large Margin On Design Limits - Does Not Support Envelope Expansion - Minimum Testing To Get Aircraft Flying Altitude ~50% V L Design V L Mach Number Detailed GVT After Variant 1 st Flight ~50% M L Incremental Flight Envelope Clearance Expansion As Flutter Testing Progresses Design M L - Predicted Flutter Margin At Design Speeds Exceeds Requirements Altitude - Predicted Buffet Response Is Within Design Limit Loads - Supports Envelope Expansion Testing ~50% V L Design V L - Supports Clearance To Design Limits Mach Number 20
Future Use Of Verified FE Models Verification Tests Provide Confidence In Finite Element Models Aeroelastic Stability Analyses Using These Models Likewise Increases In Confidence Analysis Will Be Used To Virtually Demonstrate That The F-35 Meets All Aeroelastic Stability Margins That Cannot Be Directly Demonstrated On The Aircraft Flutter And Divergence Airspeed Margins Aeroservoelastic (ASE) Gain And Phase Margins Streamlined Future Structural Dynamics Certification Activity Improved Reliability For Targeting Critical Configurations Reduced Ground And Flight Testing Requirements Highly Representative FEM Can Be Used To Investigate Future Structural Modifications Or Changes With Realistic FEM Component Modifications Yields High Confidence Aeroelastic Assessment Before Commencing The Flight Test Verification Program 21
Questions? 22