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mostly constant position on the aft portion of the wing, parallel to the TE. Differences are expected in the tip-launcher region, such as the small secondary shock aft of the primary shock near the tip launcher due to the tip launcher notch seen in the roll oscillation case. The most interesting difference, however, is aft of the shock on the inboard portion of the wing where apparent separation of the shock is evident. This is much more visible upon animation. For the clean-wing case, this separation appears to occur all the way out to the wing tip and be much more significant overall. For the tip-launcher case, it looks like the separation begins to occur near the tip, but then interacts with the secondary shock due to the launcher, resulting in less separation. This same trend is seen at the inboard separation. Upon animation of Figure 6-39 C) and D), the clean-wing case reveals that as the aircraft pitches nose-down, the Mach=1 iso-surface wraps around the wingtip at the LE extending about one-half the length of the wing. This surface then wraps around the bottom surface of the wing at the LE, and continues along the LE, tapering off well inboard. Upon animation of the tip- launcher case in Figure 6-39 A) and B), it is revealed that as the aircraft pitches nose-down, the Mach=1 iso-surface is inhibited from wrapping around the wingtip at the LE by the tip launcher. It wraps around the tip launcher at about one-half the way down the tip. The shock along the LE still wraps around the bottom surface of the wing. For both cases, the presence of LE antenna causes the shock to snake along the bottom surface of the wing and connect to the base of the LE antenna just like for the roll case. However, it also appears that the forward shock on the LE antenna attaches to the shock on the LE edge of the wing. From animation of Figure 6-40 A) and B), vortices are seen rotating along the tip launcher, corresponding to the pitching up and down of the aircraft. Rotation in the LE antenna vortex is also apparent as the aircraft is pitching up and down. The tip vortex and LE antenna vortex share 94
the same direction of rotation on the upper surface, but are contra on the lower surface. The LE antenna vortex diminishes on the upper surface as the aircraft pitches down and grows as it pitches up. Additionally, there is much more activity from the strake vortex on inboard portion of the wing. This feature may be contributing to the inboard separation of the shock seen previously. This may be strong evidence of shock-induced separation. Additionally, this type of flow feature may play an important role in the LCO mechanism. The instantaneous Cp measurements plotted against non-dimensional chord are plotted in Figure 6-41 at 98% span, in Figure 6-42 at 93% span, and in Figure 6-43 at 88% span on the left wing (looking aft- forward) for one developed cycle of oscillation. The 88% span location is chosen as the region of interest due to its vicinity to the underwing missile launcher location and will be discussed most extensively. Results for the Gird8 tip-launcher case are shown in A), upper surface, and B), lower surface; and the Grid9 clean-wing comparison results are illustrated in C), upper surface, and D), lower surface. Each numbered line (1-9) corresponds to the time- accurate oscillation cycle angle, in degrees, during the sinusoidal pitching cycle, indicated by the corresponding cycle angle color given in the legend in the upper right-hand corner. In Figure 6-43, the Cp on both the upper and lower surface overlap for 0° and 360° cycle angles, but do not for 180°. Similar to the clean-wing pitching case, from a quasi-static perspective it can be expected that the coefficients would be the same for all of these cycle angles since they are at the same absolute AOA. Upon comparison of the tip-launcher case and the clean-wing case, the same cycle angles create pairs, indicating the same inability of the flow- field to “keep up” with the motion of the structure, indicating a lag in the flow. The same suction loss feature is observed for cycle angles 225°, 270°, and 315°, corresponding to the ascension, peak, and descension of the positive peak of the pitch oscillation in the 65-75% chord region on 95
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TEMPORAL ANALYSIS OF TRANSONIC FLOW
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To Dodjie 3
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TABLE OF CONTENTS ACKNOWLEDGMENTS .
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Summary ...........................
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LIST OF FIGURES Figure page 2-1 F-1
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6-9 Lissajous plots of upper surfac
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6-45 2-D (left column) and 3-D (rig
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incrementally since this capability
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are not sufficient to provide an an
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4. Contribute to the work of others
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Classical aircraft flutter is chara
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Popular methods for finding the sol
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LCO amplitude but not the mechanism
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speed for the particular configurat
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Denegri 23 explains the flutter fli
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Doublet-lattice Denegri 1, 7,12,13
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Batina’s work and uses an F-16 ha
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As a result of Denegri’s previous
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Figure 2-1. F-16 store stations. Fi
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Figure 2-5. Flutter flight test bui
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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
Page 67 and 68: Figure 5-11. CFD vs. wind tunnel co
Page 69 and 70: Figure 5-13. CFD vs. wind tunnel co
Page 71 and 72: server cluster comprised of 5,120 p
Page 73 and 74: of oscillation. The 88% span locati
Page 75 and 76: since it is proportional to lift, a
Page 77 and 78: of phase. Figure 6-10 B) and D) are
Page 79 and 80: The instantaneous Cp measurements p
Page 81 and 82: D Lissajous plots with time being t
Page 83 and 84: frequency of 8Hz. The wavelet plot
Page 85 and 86: vortex is observed travelling outbo
Page 87 and 88: near 70% chord and the second near
Page 89 and 90: expected from a lift curve slope. A
Page 91 and 92: shape of this Lissajous is nonlinea
Page 93: occurring at harmonics, and assumes
Page 97 and 98: the roll cycle, and the black arrow
Page 99 and 100: B) shows the fast Fourier transform
Page 101 and 102: is that shock separation aft of the
Page 103 and 104: on the upper surface, a much larger
Page 105 and 106: Figure 6-1. F-16 wing planform with
Page 107 and 108: A C B Figure 6-4. DDES of F-16 in s
Page 109 and 110: Cp (Non-dimensional) Cp (Non-dimens
Page 111 and 112: -Cp (Non-dimensional) A -Cp (Non-di
Page 115 and 116: -Cp (Non-dimensional) A Time (sec)
Page 117 and 118: A B Figure 6-14. DDES-SARC of F-16
Page 121 and 122: Figure 6-18. Lissajous plots of upp
Page 123 and 124: -Cp (Non-dimensional) Frequency (Hz
Page 125 and 126: -Cp (Non-dimensional) A Frequency (
Page 127 and 128: A B Figure 6-24. DDES-SARC of F-16
Page 143 and 144: A B D Figure 6-40. DDES of F-16 in
Page 145 and 146:
Cp (Non-dimensional) Cp (Non-dimens
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Figure 6-44. Lissajous plots of upp
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-Cp (Non-dimensional) A Time (sec)
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-Cp (Non-dimensional) A Time (sec)
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A B Figure 6-50. DDES-SARC of F-16
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Figure 6-52. Lissajous plots of upp
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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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