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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, 22-25 June 2009
Copyright © 2009 by Lockheed Martin Corporation
A08-23854013
Unsteady aero can be destructive
Buffet
Wake vortex encounter
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 Flight Experiences Wind Tunnel Testing Observations
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 Flight Experiences Wind Tunnel Testing Observations
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
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 Flight Experiences Wind Tunnel Testing Observations
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
A03-08521011
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
A03-08521012
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
F-16 Ventral Fin Buffet (Concluded)
A03-08521013
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
A03-08521014
F-111 TACT Wing Buffet (Cont’d)
Predicted and Measured RMSAccelerations Versus Angle of Attack
A03-08521015
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
A03-08521016
F-111 TACT Wing LCO
A03-08521017
= 9.1°
= 10.0° = 11.1°
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
A03-08521018
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
A03-08521019
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
A03-08521020
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 Flight Experiences Wind Tunnel Testing Observations
Model Geometry and Instrumentation/Flow-Visualization Locations
Force/Pressure Model Geometry and Instrumentation Layout
Flow-Visualization Light Sheet Positions
108.65
76°
AA
5
Section A-A
0.73
3 C
roo
t
820.
7Pitching Axis
Bearings
Shaft6
7Wing
Balance
4 3 2 1
62.3 7.0
Hydraulic Actuator
Wind Tunnel Sidewall
TurntableAccelerometersPressure Section
A05-15010003
Conf. 5 Conf. 2 Conf. 1
9
8
7
RotationAxis
14
1
2
3
4
5
6
10
11
12
13(T.E.)
Sheet Angle From Vertical = 4.7 deg
Flow-Visualization Technique andImage Orientation
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
High Speed Video CameraOptical System (OS) (on Travel Mechanism)
Water Input
Camera
HST Side View
ModelFlow
Mirror Slat With Observation Widows
XZ
HSVC OS
Laser Light Sheet
Pitching Axis
Side Camera
FlowHST Cross-Section
Pressure Row
4 3 2 1
Trailing Edge
Wing Leading
Edge
Targets
Strake
A05-15010004
Steady Pressure Distributions on the Clean Wing
• M = 0.90, AOA = 6.45 deg to 11.39 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
-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
ALPHA = 10.39 deg
ALPHA = 11.39 deg
Attached Attached, Pre-SITES
SITES Leading-Edge Separation
A05-15010006
Pulsed Laser Flow-Visualization, Transition From Attached to Leading-Edge Separation
Pulsed Laser Sheet at 80% Span Oriented Streamwise 9 Nano-Sec Pulse Duration
A05-15010007
Attached Flow Normal Shock Small Separation Bubble
at Foot of Shock
SITES “LAMBDA” Shock Larger Region of Shock-
Induced Separation
Leading-Edge Separation
ALPHA = 10.0 DEG ALPHA = 10.5 DEG ALPHA = 11.0 DEGLeading Edge Trailing Edge Leading Edge Trailing Edge Leading Edge Trailing Edge
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
A05-15010008
Comparison of Clean Wing and Tip Missile Configurations, AOA = 8.5 Deg
M = 0.9, Spanwise Sheet Positions 11, 12, 13
A05-15010010
AOA = 8.5°
Comparison of Clean Wing and Tip Missile Configurations, AOA = 10.5 Deg
Mach = 0.9, Spanwise Sheet Positions 11, 12, 13
A05-15010011
AOA = 10.5°
Natural Unsteadiness for Clean Wing and Tip Missile Configurations
M = 0.9, Sheet Positions 12 and 13 - 640 Frames/sec! Transition From SITES to Leading-Edge Separation
A05-15010012
Clean Wing, Sheet Position 12, AOA = 11 Deg
Wing With Tip Missile, Sheet Position 13, AOA = 9.51 Deg
Frame = 871 Frame = 873 Frame = 875 Frame = 877
Frame = 879 Frame = 880 Frame = 887 Frame = 888
Oscillatory Pitch Motion for the Tip Missile Configuration, AOA = 8.0 Deg
M = 0.9, Amplitude = ±0.5 Deg, Frequency = 40 Hz
AOARange
A05-15010013
MaximumAOA = 8.5°
MinimumAOA = 7.5°
Oscillatory Pitch Motion for the Tip Missile Configuration, AOA = 9.0 Deg
A05-15010014
M = 0.9, Amplitude = ±0.5 Deg, Frequency = 40 Hz
AOARange
MaximumAOA = 9.5°
MinimumAOA = 8.5°
Oscillatory Pitch Motion for the Tip Missile Configuration, AOA = 10.0 Deg
M = 0.9, Amplitude = ±0.5 Deg, Frequency = 40 Hz
A05-15010015
MaximumAOA = 10.5°
MinimumAOA = 9.5°
AOARange
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 Flight Experiences Wind Tunnel Testing Observations
High Rate Maneuvers and Other Transients
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
During rapid maneuvers, the aircraft kinematic state (accels, rates, etc. may not be indicative of the aircraft loads state
High Rate Maneuvers and Other Transients
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
During rapid maneuvers, the aircraft kinematic state (accels, rates, etc. may not be indicative of the aircraft loads state
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 Flight Experiences Wind Tunnel Testing Observations
High Rate Transients
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
Time lags in flow state transitions and control law changes are problematic
A08-23854019
1.5
1.
0.5
0.
-0.5-10. 0. 10 20. 30. 40.
CNCN
DEG DEG
2.
50.
1.5
1.
0.5
0.
-0.5
-10. 0. 10 20. 30. 40.
DEG
2.
50.
CNCN
1.5
1.
0.5
0.
-0.5
-10. 0. 10 20. 30. 40.
DEG DEG
2.
50.
1.CNCN
High Rate Transients
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
Time lags in flow state transitions and control law changes are problematic
A08-23854004
g
Roll Rate (deg/s)
10 20 30 40
8º
LEF=7º
9º
10º
7º
High-g Maneuvers
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
Static aeroelasticity can be highly non-linear where wing tip flow separation is present
CFD based aeroelastic solution for an F-16 high-g maneuver
A08-23854020
Wake Vortex Encounters
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
These can range from annoying bumps to structural damage to loss of aircraft
A08-23854012
S-80 encounter in DFWLanding path ~ 5K FT
Left wing uplift, right wing downRWD roll back to neutral
Left wing down, right wing uplift LWD roll
Left wing upliftRWD roll
A08-23854013
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 Flight Experiences Wind Tunnel Testing Observations
Model Geometry and Instrumentation/Flow-Visualization Locations
Force/Pressure Model Geometry and Instrumentation Layout
Flow-Visualization Light Sheet Positions
108.65
76°
AA
5
Section A-A
0.73
3 C
roo
t
820.
7Pitching Axis
Bearings
Shaft6
7Wing
Balance
4 3 2 1
62.3 7.0
Hydraulic Actuator
Wind Tunnel Sidewall
TurntableAccelerometersPressure Section
A05-15010003
Conf. 5 Conf. 2 Conf. 1
9
8
7
RotationAxis
14
1
2
3
4
5
6
10
11
12
13(T.E.)
Sheet Angle From Vertical = 4.7 deg
Mach Effects on Steady Normal Forceand Pitching Moment
M = 0.225, 0.6, and 0.9
1.2
0.8
0.4
CN
0.12
0.08
0.04
CM
0
00 10 20 30 40 50
0.225 8.0
0.60 8.0
0.90 8.0/15.0
M Re x 10-6
A01- 01100015
Rapid High AOA Maneuvers with Sideslip for Low Speed Full Span Straked Wing Model
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
Rapid post stall maneuver airloads are highly path dependent
A08-23854018
ClCl
DEG DEG
0.02
0.01
0.
- 0.01
- 0.02
- 0.03
- 0.04
- 0.05
- 0.06-10. 0. 10 20. 30. 40. 50.
DEG DEG
0.02
0.01
0.
- 0.01
- 0.02
- 0.03
- 0.04
- 0.05
- 0.06-10. 0. 10 20. 30. 40. 50.
A08-23854018
ClCl
DEG DEG
0.02
0.01
0.
- 0.01
- 0.02
- 0.03
- 0.04
- 0.05
- 0.06-10. 0. 10 20. 30. 40. 50.
DEG DEG
0.02
0.01
0.
- 0.01
- 0.02
- 0.03
- 0.04
- 0.05
- 0.06-10. 0. 10 20. 30. 40. 50.
Low pitch rate
High pitch rate
Ro
llin
g M
om
en
tR
oll
ing
Mo
me
nt
Attached Transonic Flow, AOA ≈ 9 Deg
Outer Panel Forward and Aft Normal Shocks Beginning of Strake Vortex Flow
A01-01100003
Leading-Edge Separation of Outboard Panel, AOA ≈ 11.5 Deg
Outboard of Section 1 Continued Development of Strake Vortex
A01- 01100005
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
A01- 01100004
Leading-Edge Separated Flows atAOA = 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)
A01- 01100007
Transonic Vortex Flow, AOA ≈ 12 Deg to 19 Deg
Strake Vortex Dominant Shock Observed in Vortex Core
A01- 01100006
Shocklet/Finger Vortex Flow,AOA ≈ 19 Deg to 27 Deg
Strake Vortex Bursting Multiple Shocks and Vortex Pairs Existing Above Separated Flows
A01- 01100008
Transition to Turbulent Separation BoundaryFlow, AOA ≈ 27 Deg
Progressive Stalling Vortex Bursting, Inboard Wing Panel Lifting
Shocks and Vortices above Separated Flows Vortex Bursting, Outboard Wing Panel Still Lifting
A01- 01100008
Progressive Stall Development, AOA >27 Deg
Progressive Stalling Strake Vortex Bursting
Progressive Stalling Wing Panel Completely Stalled Vortex Bursting Seen on Strake
A01- 01100012
Unsteady Normal Force and PitchingMoment at M = 0.9
Pitch Oscillation From = 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
0
0.08
0.12
SITES and Wing Tip Separation
“Conventional” Vortex Breakdown and Stalling
M = 0.9
0 10 20 30 40 50
Dynamic, Pitch-Up
Dynamic, Pitch-Down
Static
A01- 01100017
Unsteady Flow for Oscillatory PitchingBetween 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
A01- 01100018
Unsteady Flow for Oscillatory PitchingBetween 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
A01- 01100019
Unsteady Flow for Oscillatory PitchingBetween 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
A01- 01100020
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
A01-01100021
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
A01-01100021
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
A01-01100021
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.
A01-01100024