Real Life Adventures with Unsteady Aerodynamics by Dr. Atlee M

Name: Real Life Adventures with Unsteady Aerodynamics by Dr. Atlee M
Uploaded: 2017-09-08T04:55:46+00:00
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Description: Real Life Adventures with Unsteady Aerodynamics by Dr. Atlee M

Real Life Adventures with Unsteady Aerodynamics by Dr. Atlee M
Real Life Adventures with Unsteady Aerodynamics by Dr. Atlee M. Cunningham, Jr. Lockheed Martin Senior Fellow Lockheed Martin Aeronautics Company, Fort Worth, Texas Presented for “Aerodynamic and Fluid Dynamic Challenges in Flight Mechanics” Working Group Meeting 27th AIAA Applied Aerodynamics Conference San Antonio, Texas, June 2009 Unsteady aero can be destructive Buffet Wake vortex encounter Copyright © 2009 by Lockheed Martin Corporation

The Influence of Unsteady Aerodynamics is Multi-Faceted
Design, development, and testing of today’s aircraft involves the close integration of many diverse technologies and systems which is expected to become more complex in the future. For example, these technologies cover Airframe – aerodynamics, structures, aeroelasticity, aeroservoelasticity (ASE) Flight controls – active controls, g/AOA limiters, weapons delivery Propulsion – thrust management/vectoring, auxiliary systems Stores/weapons – internal/external carriage, delivery systems Maintainability – logistics, spares, inspections, service life Pilot/crew – comfort, performance, limits, communications

The Influence of Unsteady Aerodynamics is Multi-Faceted – More So for Fighter Aircraft
Fighter aircraft can be more complex than transport aircraft due to: Multi-role – air-to-air, air-to-ground, large stores inventory Highly transient maneuvers – offensive/defensive, weapons release Operations at edge of envelope – high AOA, transonic and buffeting conditions Susceptibility to wake vortex encounters – air-to-air tail chasing Fighter design envelopes contain many loads conditions that are not steady state: Rapid maneuvers – unsteady aerodynamics, changing control laws, buffeting Edge of the envelope – non-linear, path dependency, extreme conditions Complex systems interactions – fail-safe designs, redundancies

Real Life Adventures with Unsteady Aerodynamics Outline
Aircraft Buffeting and Limit Cycle Oscillations (LCO) Flight Experiences Wind Tunnel Testing Observations High Rate Maneuvers and Other Transients

Differences Between Buffeting and LCO
Structural buffet response is driven by unsteady separated flows acting on the structure Unsteady flows unaffected by structural motion Forcing is wide band and affects many structural modes of vibration LCO is a nonlinear interaction between aerodynamics forces and structural response Similar to flutter but limited in response amplitude Nonlinear forces act to drive the LCO

Why are buffeting and LCO important
Buffeting sources Spoilers and deployed flaps during landing Wing mounted stores and protuberances Weapons and landing gear bays Leading edge separation and vortex breakdowns Shock induced separation Inlet spillage and thrust reversers Buffeting problems Fatigue of TE controls, spoilers, etc. Fatigue of fins, antennae, twin vertical tails and other downstream surfaces Severe vibrations of wires/cables/equipment/bulkheads in open bays

Why are buffeting and LCO important (cont’d)
LCO sources Transonic speeds, freeplay, nonlinear damping Embedded shocks and induced flow separation Susceptible structural vibrations with low damping (sensitive to various nonlinearities) LCO problems False indications of flutter onset Pilot discomfort/distractions/limitations Weapons system limitations Control surface buzz

F-16 Ventral Fin Buffet Early F-16 Ventral Fins Were Subject to Partial or Complete Loss During Flight Inlet Spillage Turbulence Was Primary Buffeting Source Safety of Flight Issues Were Marginal Biggest Problems Were With Maintenance and Spares Requirements to Carry LANTIRN Pods Upstream of the Ventrals Were Imposed in the Early 1980’s First Flight With LANTIRN’S Resulted in Loss of R.H. Ventral Fin Redesign of Ventrals and Supporting Structure Was Accomplished but Was Not Good Enough Later Structural Redesign Was Also Incorporated but Problems Still Existed With Cracking and Loss of Fins A

F-16 Ventral Fin Buffet (Cont’d)
M=0.95, 10 Kft, Clean Aircraft With/Without LANTIRN Pods, 1-g Level Flight LANTIRN Pods’ Wake Turbulence Almost as Severe as Inlet Lip Spillage During Throttle Chop High-Amplitude Responses Exist Continuously Even at Lower Machs LANTIRN’S Somewhat Reduce Throttle Chop Effects A

F-16 Ventral Fin Buffet (Concluded)
An Investigation Was Conducted in the Mid-1990’s To Redesign the Ventrals and Supporting Structure Tests Conducted in Fort Worth (Upgraded Block 40 F-16) and The Netherlands (Early Block 15 F-16) Aerodynamics and Structural Modifications Evaluated Four Fin Configurations Tested Baseline Block 40 (BSLN) Block 40 With Stiffer Skins (MMC Aluminum) (MMC) Block 40 With Stiffer Skins and an Aerodynamic Nose Cap (MMC NC) Thicker Fin With an Airfoil Section Shape (NACA) The MMCNC Fin Has Been Adopted as the Only Spare Fin for All F-16’s Failure Rates Have Dropped Dramatically A

F-111 TACT Wing Buffet An F-111 Was Used as a Test Bed for Investigating the Potential Benefits of Transonic AirCraft Technology (TACT) Conducted by NASA, AFRL and General Dynamics Fitted With a Supercritical Wing With Variable Sweep Extensive Flight Test Program Highly Instrumented Wing for Buffet Research 1/6-Scale Wind Tunnel Model (Steel and Aluminum Wings, Instrumented With Pressure Transducers in Same Locations as on the A/C) Buffet Prediction Research Was Conducted by NASA Ames and General Dynamics Used Pressure Time Histories With Mode Deflections To Obtain Generalized Force Time Histories Predictions Made With Equations of Motion and Aerodynamic Damping Derived From the Wind Tunnel Model Response A

F-111 TACT Wing Buffet (Cont’d)
Predicted and Measured RMS Accelerations Versus Angle of Attack A

F-111 TACT Wing Buffet (Concluded)
Buffet Predictions for the Wing Bending Mode Agreed Well With Flight Test Data AOA Range From 7 Deg to 12 Deg, M=0.8 Wing Sweeps of 26 Deg and 35 Deg Torsion Mode Predictions Were Mixed Good Agreement for Wing Sweep of 35 Deg Significant Under Prediction for 26 Deg Sweep Similar Results From Earlier Buffet Research for the F-111 About the Same AOA, Wing Sweep and Mach Number Ranges Strong Suspicion That a Torsion Mode LCO Existed A

F-111 TACT Wing LCO a = 9.1° a = 10.0° a = 11.1° A

F-111 TACT Wing LCO (Concluded)
A Step Change in Pitching Moment With the Onset of Shock-Induced Trailing Edge Separation Was Key to Driving a Torsion Mode LCO Step Increase in Nose-Down Pitching Moment With Increasing AOA Aerodynamic Lag Produces an Unstable Hysteresis Loop Aerodynamic and Structural Damping Counteract the Unstable Loop A Simple One DOF Math Model for the Torsion Mode With a Nonlinear Step Change in Generalized Force Produced an LCO Time History Solution to the One DOF Equation of Motion A

F-16 Wing LCO F-16 Wing LCO Was Reported in the Early 1980’s for Certain Air-to-Air Wing Store Combinations During Wind-Up-Turns AOA in the 5 Deg to 7 Deg Range, M=0.90 to 0.96 Conditions Corresponded to Onset of Shock-Induced Tailing Edge Separation (Similar to F-111 TACT) An Investigation Was Funded by the USAF to Investigate the Phenomenon Cooperative Program Between General Dynamics and the National Aerospace Laboratory in The Netherlands Extensive Wind Tunnel Tests at Transonic Speed With About 100 High Response Pressure Transducers on the Wing Oscillating Wing Panel Analytical Studies To Develop Prediction Methodology “NONLINAE” Was the Result A

F-16 Wing LCO (Concluded)
Predictions of Wing LCO for the Early F-16 Version Agreed Well With Flight Test Data Strong Sensitivity to Structural Damping Suggested Probable Source of A/C to A/C Variations in LCO Levels Minimum of 2 Critical Modes to Reproduce LCO Problem Was Significantly Reduced With Subsequent Structural Upgrades Stiffer Wing Was Developed for Block 40 F-16s To Accommodate Higher Gross Weights A

Model Geometry and Instrumentation/ Flow-Visualization Locations
108.65 76° 0.733 Croot 820.7 Sheet Angle From Vertical = 4.7 deg A A Section A-A 7 5 Bearings 14 6 5 1 6 Shaft 2 8 Pitching Axis 3 10 Rotation Axis 4 7 Wing Balance 11 9 12 4 3 2 1 Hydraulic Actuator 62.3 7.0 13 (T.E.) Accelerometers Pressure Section Turntable Conf. 5 Conf. 2 Conf. 1 Wind Tunnel Sidewall Force/Pressure Model Geometry and Instrumentation Layout Flow-Visualization Light Sheet Positions A

Flow-Visualization Technique and Image Orientation
Pressure Row Optical System (OS) (on Travel Mechanism) High Speed Video Camera Mirror Slat With Observation Widows Water Input Z X Camera Model Flow Trailing Edge HST Side View HSVC OS Laser Light Sheet Side Camera Wing Leading Edge Pitching Axis Flow HST Cross-Section Targets Strake Set-Up for Spanwise Sheet Visualization Sheet Rotated 90° for Chordwise Sheet Use Side Camera Pulsed Laser Model Relationship With Spanwise Sheet at “Position 9” Image Reversed (Negative) for Improved Quality A

Steady Pressure Distributions on the Clean Wing
M = 0.90, AOA = 6.45 deg to deg -1.0 -0.8 -0.6 -0.4 -0.2 0.0 -1.0 -0.8 -0.6 -0.4 -0.2 0.0 ALPHA = 6.45 deg ALPHA = 8.39 deg Attached Attached, Pre-SITES ALPHA = deg ALPHA = deg -1.0 -0.8 -0.6 -0.4 -0.2 0.0 -1.0 -0.8 -0.6 -0.4 -0.2 0.0 SITES Leading-Edge Separation A

Pulsed Laser Flow-Visualization, Transition From Attached to Leading-Edge Separation
Pulsed Laser Sheet at 80% Span Oriented Streamwise 9 Nano-Sec Pulse Duration ALPHA = 10.0 DEG ALPHA = 10.5 DEG ALPHA = 11.0 DEG Leading Edge Trailing Edge Leading Edge Trailing Edge Leading Edge Trailing Edge Attached Flow Normal Shock Small Separation Bubble at Foot of Shock SITES “LAMBDA” Shock Larger Region of Shock-Induced Separation Leading-Edge Separation A

Spanwise Location of Separation for Tip Missile/Launcher Configuration
M = 0.9, AOA = 6 deg to 10.5 deg, Spanwise Sheet Positions 11, 12, 13 Attached Flow Up to 8.0 deg Transition to SITES at 8.0 deg, Continues up to 9.5 deg Not Full Chordwise Separation Leading-Edge Separation Transition at 10.5 deg Full Chordwise Separation A

Comparison of Clean Wing and Tip Missile Configurations, AOA = 8.5 Deg
M = 0.9, Spanwise Sheet Positions 11, 12, 13 AOA = 8.5° A

Comparison of Clean Wing and Tip Missile Configurations, AOA = 10
Comparison of Clean Wing and Tip Missile Configurations, AOA = 10.5 Deg Mach = 0.9, Spanwise Sheet Positions 11, 12, 13 AOA = 10.5° A

Natural Unsteadiness for Clean Wing and Tip Missile Configurations
M = 0.9, Sheet Positions 12 and Frames/sec! Transition From SITES to Leading-Edge Separation Frame = 871 Frame = 873 Frame = 875 Frame = 877 Clean Wing, Sheet Position 12, AOA = 11 Deg Frame = 879 Frame = 880 Frame = 887 Frame = 888 Wing With Tip Missile, Sheet Position 13, AOA = 9.51 Deg A

Oscillatory Pitch Motion for the Tip Missile Configuration, AOA = 8
Oscillatory Pitch Motion for the Tip Missile Configuration, AOA = 8.0 Deg M = 0.9, Amplitude = ±0.5 Deg, Frequency = 40 Hz Maximum AOA = 8.5° Minimum AOA = 7.5° AOA Range A

High Rate Maneuvers and Other Transients
During rapid maneuvers, the aircraft kinematic state (accels, rates, etc. may not be indicative of the aircraft loads state Path dependency is very important Flow separation induced time lags are significant Lag for re-attaching flow is higher than separating flow “Static” flex-to-rigid ratios are different for attached or separated flows Conventional ASE analysis tools do not account for these effects System induced time lags and false state conditions can be destabilizing Data acquisition, processing and transferring to commands depends on sample rates and sensors Controller/actuator lags are more significant Structural vibration modes can introduce false state conditions

High Rate Maneuvers and Other Transients
During rapid maneuvers, the aircraft kinematic state (accels, rates, etc. may not be indicative of the aircraft loads state Store ejection and wake vortex encounters are very sensitive to the current aircraft conditions Rapid control law changes due to store downloads cannot be assessed with current ASE methods and may be critical if active flutter suppression is used Aircraft response to wake vortex encounters is also affected by pilot/control system commands and how the wake is entered

High Rate Transients Flow separation State change is very quick
Time lags in flow state transitions and control law changes are problematic Flow separation State change is very quick Lag for flow separating is much less than for reattachment Leading/trailing edge and shock induced separation are affected Max lift overshoot on pitch-up can occur Control law changes Data processing and structural dynamics can induce lagging Actuator response is more lagging

High Rate Transients High-g roll maneuver anomaly
Time lags in flow state transitions and control law changes are problematic High-g roll maneuver anomaly RWD roll initiated during high-g symmetric maneuver Right wing tip flow probably separated Sudden re-attachment on right wingtip reduced roll rate and increased g’s Attributed to LE flap change from 7° to 10° during maneuver

CFD based aeroelastic solution for an F-16 high-g maneuver
High-g Maneuvers Static aeroelasticity can be highly non-linear where wing tip flow separation is present Spanwise bending moments are less if the wing tip is separated Shock induced separation is most common Flex-to-rigid ratios are higher Wing washout at high-g’s and transonic speeds can eliminate shock induced separation Nose-down twist from wash-out weakens tip shocks Verified by CFD solutions for a wind tunnel model investigation CFD based aeroelastic solution for an F-16 in a high-g maneuver demonstrated this effect Large wing tip deflections of over 20 in. Weakened wing tip shocks Correlated well with flight test data CFD based aeroelastic solution for an F-16 high-g maneuver

Wake Vortex Encounters
These can range from annoying bumps to structural damage to loss of aircraft Loss of Airbus A300, AA Flt 587, 17 Nov 2001 A300 followed 747 on climb-out Loss attributed to pilot over-reacting to severe wake turbulence Uncommanded double roll on approach to DFW Occurred at about 5k ft in landing pattern Rapid double roll in about 1 second No post encounter rolling Losses of F-16 ventral tail tips Occurs during air-to-air combat training exercises (tail chasing) Four incidents since 1980’s Attributed to wing tip vortex from lead aircraft Pilot unaware of loss

Model Geometry and Instrumentation/ Flow-Visualization Locations
108.65 76° 0.733 Croot 820.7 Sheet Angle From Vertical = 4.7 deg A A Section A-A 7 5 Bearings 14 6 5 1 6 Shaft 2 8 Pitching Axis 3 10 Rotation Axis 4 7 Wing Balance 11 9 12 4 3 2 1 Hydraulic Actuator 62.3 7.0 13 (T.E.) Accelerometers Pressure Section Turntable Conf. 5 Conf. 2 Conf. 1 Wind Tunnel Sidewall Force/Pressure Model Geometry and Instrumentation Layout Flow-Visualization Light Sheet Positions A

Mach Effects on Steady Normal Force and Pitching Moment
M = 0.225, 0.6, and 0.9 1.2 M Re x 10-6 CN 0.8 /15.0 0.4 0.12 CM 0.08 0.04 10 20 30 40 50 a A

Rapid High AOA Maneuvers with Sideslip for Low Speed Full Span Straked Wing Model
Rapid post stall maneuver airloads are highly path dependent Symmetric pitch maneuvers can produce max lift overshoots Buffet on pitch-up is low Buffet on pitch-down is higher and more persistent Asymmetric pitch maneuvers can affect roll moments Dynamic distortion through unstable regions is very sensitive to pitch rate Effects are very path dependent Example – pitching model with sideslip Large AOA pitching motions at -5 deg sideslip Unstable roll moments between about 16 deg to 38 deg AOA for steady flow Motions limited to minimum flow state changes distort the steady characteristics Motions that cross several flow state changes completely change the character Rolling Moment Low pitch rate Rolling Moment High pitch rate

Attached Transonic Flow, AOA ≈ 9 Deg
Outer Panel Forward and Aft Normal Shocks Beginning of Strake Vortex Flow A

Leading-Edge Separation of Outboard Panel, AOA ≈ 11.5 Deg
Outboard of Section 1 Continued Development of Strake Vortex A

Transition to Shock-Induced, Trailing-Edge Separation (SITES), AOA ≈ 10.5 Deg
Pulsed Laser Sheet at 80% Span Oriented Streamwise 9 Nano-Sec Pulse Duration Attached Flow Normal Shock Small Separation Bubble at Foot of Shock SITES “LAMBDA” Shock Larger Region of Shock-Induced Separation A

Leading-Edge Separated Flows at AOA = 11.0 Deg and 22.0 Deg
Pulsed Laser Sheet at 80% Span Oriented Streamwise 9 Nano-Sec Pulse Duration Initial Occurrence of Leading- Edge Separation Supersonic Flow Above Separation Zone Separation Shear Layer Structure Visible Fully Developed Leading-Edge Separation Supersonic Flow More Pronounced Above Separation Shear Layer Structure More Distinct (Shocklets and Finger Vortices) A

Transonic Vortex Flow, AOA ≈ 12 Deg to 19 Deg
Strake Vortex Dominant Shock Observed in Vortex Core A

Shocklet/Finger Vortex Flow, AOA ≈ 19 Deg to 27 Deg
Strake Vortex Bursting Multiple Shocks and Vortex Pairs Existing Above Separated Flows A

Transition to Turbulent Separation Boundary Flow, AOA ≈ 27 Deg
Shocks and Vortices above Separated Flows Vortex Bursting, Outboard Wing Panel Still Lifting Progressive Stalling Vortex Bursting, Inboard Wing Panel Lifting A

Progressive Stall Development, AOA >27 Deg
Progressive Stalling Strake Vortex Bursting Progressive Stalling Wing Panel Completely Stalled Vortex Bursting Seen on Strake A

Unsteady Normal Force and Pitching Moment at M = 0.9
Pitch Oscillation From a = 7 deg to 37 deg at 3.8 Hz Lagging of Flow Transitions on Both Pitch-Up and Pitch-Down 1.2 0.8 0.4 CN 0.04 CM 0.08 0.12 SITES and Wing Tip Separation “Conventional” Vortex Breakdown and Stalling M = 0.9 10 20 30 40 50 a Dynamic, Pitch-Up Dynamic, Pitch-Down Static A

Unsteady Flow for Oscillatory Pitching Between 7. 2 Deg and 37
Unsteady Flow for Oscillatory Pitching Between 7.2 Deg and 37.7 Deg at 3.8 Hz Pitch-Up/Pitch-Down Effects at 11.5 Deg Pitch-Up Shows Delay of Outboard Panel Lift Breakdown Pitch-Down Produces Delay of Outboard Flow Re-Establishment A

Unsteady Flow for Oscillatory Pitching Between 7.2 Deg and 37.7 Deg at 3.8 Hz (Cont’d) Pitch-Up/Pitch-Down Effects at 15 Deg to 18 Deg Pitch-Up Shows Delay of Inboard Flow Breakdown Movement Pitch-Down Shows Delay of Outboard Flow Re-Establishment A

Unsteady Flow for Oscillatory Pitching Between 7.2 Deg and 37.7 Deg at 3.8 Hz (Cont’d) Pitch-Up/Pitch-Down Effects at 27 Deg Pitch-Up Shows Delay of Outboard Wing Stalling Pitch-Down Shows Persistence of Outboard Wing Stalling but Rapid Development of the Strake Vortex A

Concluding Remarks – Why are unsteady aerodynamics important – Buffeting
Buffeting problems Fatigue of T.E. controls, spoilers, etc. Fatigue of fins, antennae, twin vertical tails and other downstream surfaces Severe vibrations of wires/cables/hydraulic lines as well as equipment/bulkheads/weapons in open bays A

Concluding Remarks – Why are unsteady aerodynamics important – LCO
LCO problems False indications of flutter onset Pilot discomfort/distractions/limitations Weapons systems limitations Control surface buzz A

Concluding Remarks – Why are unsteady aerodynamics important – Transient conditions
High rate maneuvers and other transient problems Severe aircraft buffet High loads overshoot and excursions beyond static loads Max loads that occur under transient conditions Wing drop and stall flutter Uncontrollable flow transitions Wake vortex encounters A

Lecture Data Sources Wind Tunnel Database Summarized in This Paper Included in the RTO Database Verification and Validation Data for Computational Unsteady Aerodynamics RTO-TR-26, AC/323 (AVT) TP/19, NATO, October 2000 Cunningham, A.M. and Geurts, E.G.M., “Transonic Pressure, Force and Flow Visualization Measurements on a Pitching Straked Delta Wing at High Alpha,” Paper No. 7, NATO/RTO AVT-072/073, May 2001. Cunningham, A.M., “Buzz, Buffet and LCO on Military Aircraft – The Aeroelastician’s Nightmares,” Presented at CEAS/AIAA/NVvL IFASD, Amsterdam, The Netherlands, 4-6 June 2003. Cunningham, A.M. and Geurts, E.G.M., “Flow Visualization Investigation of Transonic Limit Cycle Oscillation Conditions for a Fighter-Type Wing with Tip Stores,” Paper No. 28, NATO/RTO AVT-123, April 2005. Cunningham, A.M. and Holman, R.J., “Time Domain Aeroelastic Solutions – A Critical Need for Future Analytical Methods’ Developments,” Paper No. 12, NATO/RTO AVT-154, May 2008. A

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