Publications

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The flow over aircraft bays exhibits many characteristics of cavity flows, namely resonant pressures that can create high structural loading. Most studies have represented these bays as rectangular cavities; however, this simplification neglects many features of the actual flight geometry which could affect the unsteady pressure field and resulting loading in the bay. To address this shortcoming, a complex cavity geometry was developed to incorporate more realistic aircraft-bay features including shaped inlets and internal cavity variations. A parametric study of these features at Mach 1.5, 2.0, and 2.5 was conducted to identify key differences from simple rectangular cavity flows. The frequency of the basic rectangular cavity modes could be predicted by theory; however, most complex geometries shifted these frequencies. Geometric changes that constricted the flow tended to enhance cavity modes and create higher pressure fluctuations. Other features, such as a leading edge ramp, lifted the shear layer higher with respect to the aft cavity wall and led to cavity tone suppression. Complex features that introduced spanwise non-uniformity into the shear layer also led to a reduction of cavity tones, especially at the aft end of the cavity.

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Experiments were performed to understand the complex fluid-structure interactions that occur during internal store carriage. A cylindrical store was installed in a cavity having a length-to-depth ratio of 3.33 and a length-to-width ratio of 1. The Mach number ranged from 0.6 - 2.5 and the incoming turbulent boundary layer thickness was about 30-40% of the cavity depth. Fast-response pressure measurements provided aeroacoustic loading in the cavity, while triaxial accelerometers and laser Doppler vibrometry provided simultaneous store response. Despite occupying only 6% of the cavity volume, the store significantly altered the cavity acoustics. The store responded to the cavity flow at its natural structural frequencies, as previously determined with modal hammer tests, and it exhibited a directional dependence to cavity resonance. Specifically, cavity tones excited the store in the streamwise and wall-normal directions consistently, while a spanwise response was observed only occasionally. The streamwise and wall-normal responses were attributed to the known pressure gradients in these directions. Furthermore, spanwise vibrations were greater at the downstream end of the cavity, attributable to decreased levels of flow coherence near the aftwall. Collectively, the data indicate the store response to be dependent on direction of vibration and position along the length of the store.

A high-speed schlieren system was developed for the Sandia Hypersonic Wind Tunnel. Schlieren images were captured at 290 kHz and used to study the growth and breakdown of second-mode instabilities into turbulent spots on a 7° cone. At Mach 5, wave packets would intermittently occur and break down into isolated turbulent spots surrounded by an otherwise smooth, laminar boundary layer. At Mach 8, the boundary layer was dominated by second-mode instabilities which would break down into larger regions of turbulence. Second-mode waves surrounded these turbulent patches as opposed to the smooth laminar flow seen at Mach 5. Detailed pressure and thermocouple measurements were also made along the cone at Mach 5, 8 and 14, in a separate tunnel entry. These measurements give an average picture of the transition behavior that complements the intermittent behavior captured by the schlieren system. At Mach 14, the boundary-layer remained laminar so the transition process could not be studied. However, the first measurements of second-mode waves were made in HWT-14.

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High-frequency pressure sensors were used in conjunction with a high-speed schlieren system to study the growth and breakdown of boundary-layer disturbances into turbulent spots on a 7° cone in the Sandia Hypersonic Wind Tunnel. At Mach 5, intermittent low-frequency disturbances were observed in the schlieren videos. High-frequency secondmode wave packets would develop within these low-frequency disturbances and break down into isolated turbulent spots surrounded by an otherwise smooth, laminar boundary layer. Spanwise pressure measurements showed that these packets have a narrow spanwise extent before they break down. The resulting turbulent fluctuations still had a streaky structure reminiscent of the wave packets. At Mach 8, the boundary layer was dominated by secondmode instabilities that extended much further in the spanwise direction before breaking down into regions of turbulence. The amplitude of the turbulent pressure fluctuations was much lower than those within the second-mode waves. These turbulent patches were surrounded by waves as opposed to the smooth laminar flow seen at Mach 5. At Mach 14, second-mode instability wave packets were also observed. Theses waves had a much lower frequency and larger spanwise extent compared to lower Mach numbers. Only low freestream Reynolds numbers could be obtained, so these waves did not break down into turbulence.

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The longitudinal merging of wave packets and turbulent spots in a hypersonic boundary layer was studied on the nozzle wall of the Boeing/AFOSR Mach-6 Quiet Tunnel. Two pulsed glow perturbations were created in rapid succession to generate two closely spaced disturbances. The time between the perturbations was varied from run to run to simulate longitudinal merging. Preliminary results suggest that the growth of the trailing distur- bance seems to be suppressed by the presence of the leading disturbance. Conversely, the core of the leading disturbance appears unaffected by the presence of the trailing distur- bance and behaves as if isolated. This result is consistent with low-speed studies as well as DNS computations of longitudinal merging. However, the present results may be influ- enced by the perturber performance and therefore further studies of longitudinal merging are necessary to confirm the effect on the internal pressure structure of the interacting disturbances.

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High-frequency pressure-fluctuation measurements were made in AEDC Tunnel 9 at Mach 10 and the NASA Langley 15-Inch Mach 6 and 31-Inch Mach 10 tunnels. Measurements were made on a 7{sup o}-half-angle cone model. Pitot measurements of freestream pressure fluctuations were also made in Tunnel 9 and the Langley Mach-6 tunnel. For the first time, second-mode waves were measured in all of these tunnels, using 1-MHz-response pressure sensors. In Tunnel 9, second-mode waves could be seen in power spectra computed from records as short as 80 {micro}s. The second-mode wave amplitudes were observed to saturate and then begin to decrease in the Langley tunnels, indicating wave breakdown. Breakdown was estimated to occur near N {approx} 5 in the Langley Mach-10 tunnel. The unit-Reynolds-number variations in the data from Tunnel 9 were too large to see the same processes.

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