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CHAPTER 1 INTRODUCTION Motivation Limit Cycle Oscillation (LCO) is a sustained non-divergent periodic motion experienced by aircraft with configurations which often include stores. The external stores consist of any object carried on the aircraft, such as a missile, bomb, fuel tank, or pod. The primary concern with regard to LCO relates to the pilot’s ability to carry out his mission safely and effectively. While experiencing LCO, the pilot may have difficulty reading cockpit gauges and the heads-up display (HUD), and an unexpected LCO condition can lead to premature termination of the mission. There are also concerns regarding the effects of LCO on the stores and UAVs. These issues include such things as potential degradation in reliability of components, whether or not stores can be safely released during LCO, the possible effects of LCO on target acquisition for smart weapons, and the effects of LCO on weapon accuracy for unguided weapons. Therefore, the prediction of LCO behavior of fighter aircraft is critical to mission safety and effectiveness, and the accurate determination of LCO onset speed and amplitude often requires flight test verification of linear flutter models. The increasing stores certification demands for new stores and configurations, the associated flight test costs, and the accompanying risk to pilots and equipment all necessitate the development of tools to enhance computation of LCO through modeling of the aerodynamic and structural sources of LCO nonlinearities. Flutter, an instability caused by the aerodynamic forces coupling with the structural dynamics, and LCO are related phenomena. This association is due in part to the accuracy with which linear flutter models predict LCO frequencies and modal mechanisms. 1 However, since the characteristics of LCO motion are a result of nonlinear effects, flutter models do not accurately predict LCO onset speed and amplitude. Current engineering knowledge and theories 16
are not sufficient to provide an analytical means for direct prediction of LCO. The current approach relies heavily on historical experience and interpretation of traditional flutter analyses and flight tests as they may correlate to the expected LCO characteristics for the configuration of concern. There exists a significant need for a detailed understanding of the physical mechanisms involved in LCO that can lead to a unified theory and analysis methodology that is capable of predicting LCO responses. Many in the aeroelasticity community believe that LCO is indeed classical flutter at the onset, based on the strong similarity of the measured LCO deflection characteristics to linear flutter analysis computed deflections. However, the mechanism that bounds the oscillations still eludes the research community. Preliminary findings suggest that transonic aerodynamics is a potential bounding mechanism for typical LCO. Approach Overview The present work aims to identify flow-field features, such as pressure gradients, flow separation, and shock interactions, using computational fluid dynamics (CFD) analyses at known transonic LCO conditions observed in flight tests. The results are used to assess how characteristics in the flow may be influencing the occurrence of LCO. Examination of a true fluid-structure interaction (FSI) LCO case (flexible structure coupled with CFD) is considered quasi-incrementally since this capability does not yet exist in the flutter community. The first step is to perform fluid-structure reaction (FSR) simulations, examining the flow-field via CFD analyses on the F-16 wing during prescribed rigid body pitch and roll oscillations, simulating the torsional and bending nature of an LCO mechanism. Two configurations are examined, clean wing without 2 and with 3-5 tip missile launchers. As the state of CFD FSI codes progress, the wing will be forcefully deformed in bending, torsion, and complex LCO motions and pylons, launchers, and stores will be added. Through this build-up fluid-structure reaction (FSR) 17
Page 1 and 2: TEMPORAL ANALYSIS OF TRANSONIC FLOW
Page 3 and 4: To Dodjie 3
Page 5 and 6: TABLE OF CONTENTS ACKNOWLEDGMENTS .
Page 7 and 8: Summary ...........................
Page 9 and 10: LIST OF FIGURES Figure page 2-1 F-1
Page 11 and 12: 6-9 Lissajous plots of upper surfac
Page 13 and 14: 6-45 2-D (left column) and 3-D (rig
Page 15: incrementally since this capability
Page 19 and 20: 4. Contribute to the work of others
Page 21 and 22: Classical aircraft flutter is chara
Page 23 and 24: Popular methods for finding the sol
Page 25 and 26: LCO amplitude but not the mechanism
Page 27 and 28: speed for the particular configurat
Page 29 and 30: Denegri 23 explains the flutter fli
Page 31 and 32: Doublet-lattice Denegri 1, 7,12,13
Page 33 and 34: Batina’s work and uses an F-16 ha
Page 35 and 36: As a result of Denegri’s previous
Page 37 and 38: Figure 2-1. F-16 store stations. Fi
Page 39 and 40: Figure 2-5. Flutter flight test bui
Page 41 and 42: Stokes (RANS) equations on hybrid u
Page 43 and 44: CHAPTER 4 ANALYSIS TECHNIQUES Intro
Page 45 and 46: time-localization cannot be determi
Page 47 and 48: A B C D E Figure 4-3. Analysis of s
Page 49 and 50: (CFD) analyses in order to capture
Page 51 and 52: the appropriate time step for the n
Page 53 and 54: Wing1, is the wing-only (no fuselag
Page 55 and 56: similar to DES but with modificatio
Page 57 and 58: Figure 5-1. Time step comparison of
Page 59 and 60: A B Figure 5-3. Flow results for 8
Page 61 and 62: A B Figure 5-5. Flow results for 8
Page 63 and 64: Figure 5-7. Turbulence model compar
Page 65 and 66: Figure 5-9. CFD vs. wind tunnel com
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Figure 5-11. CFD vs. wind tunnel co
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Figure 5-13. CFD vs. wind tunnel co
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server cluster comprised of 5,120 p
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of oscillation. The 88% span locati
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since it is proportional to lift, a
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of phase. Figure 6-10 B) and D) are
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The instantaneous Cp measurements p
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D Lissajous plots with time being t
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frequency of 8Hz. The wavelet plot
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vortex is observed travelling outbo
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near 70% chord and the second near
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expected from a lift curve slope. A
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shape of this Lissajous is nonlinea
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occurring at harmonics, and assumes
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the same direction of rotation on t
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the roll cycle, and the black arrow
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B) shows the fast Fourier transform
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is that shock separation aft of the
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on the upper surface, a much larger
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Figure 6-1. F-16 wing planform with
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A C B Figure 6-4. DDES of F-16 in s
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Cp (Non-dimensional) Cp (Non-dimens
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-Cp (Non-dimensional) A -Cp (Non-di
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-Cp (Non-dimensional) A Time (sec)
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A B Figure 6-14. DDES-SARC of F-16
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Figure 6-18. Lissajous plots of upp
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-Cp (Non-dimensional) Frequency (Hz
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-Cp (Non-dimensional) A Frequency (
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A B D Figure 6-40. DDES of F-16 in
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Impact on Future Design of Aircraft
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The size of the Mach=1 iso-surface
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these cases in order to identify th
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15. Dawson, K. S., and Sussinham, J
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41. Thomas, J. P., Dowell, E. H., a
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67. Cunningham, A. M., Jr., and den
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93. Parker, G. H., Maple, R. C., an
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121. Travin, A. K., Shur, M. L., Sp
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