Fluid–structure interactions were measured between a representative control surface and the hypersonic flow deflected by it. The control surface is simplified as a spanwise finite ramp placed on a longitudinal slice of a cone. The front surface of the ramp contains a thin panel designed to respond to the unsteady fluid loading arising from the shock-wave/boundary-layer interactions. Experiments were conducted at Mach 5 and Mach 8 with ramps of different angles. High-speed schlieren captured the unsteady flow dynamics and accelerometers behind the thin panel measured its structural response. Panel vibrations were dominated by natural modes that were excited by the broadband aerodynamic fluctuations arising in the flowfield. However, increased structural response was observed in two distinct flow regimes: 1) attached or small separation interactions, where the transitional regime induced the strongest panel fluctuations. This was in agreement with the observation of increased convective undulations or bulges in the separation shock generated by the passage of turbulent spots, and 2) large separated interactions, where shear layer flapping in the laminar regime produced strong panel response at the flapping frequency. In addition, panel heating during the experiment caused a downward shift in its natural mode frequencies.
This work applies Filtered Rayleigh Scattering (FRS) for the study of shock wave/boundary layer interactions on a cone-slice-ramp geometry. As FRS measures a planar slice of the flow, the three-dimensionality of this geometry can be captured, rather than in path-integrated imaging such as schlieren. A carbon dioxide seeding system designed for the Sandia Hypersonic Wind Tunnel provides sufficient light scattering for FRS measurements. Strong background rejection in the images was achieved using a molecular gas filter, resulting in detailed visualization of flow structures within the boundary and shear layers. Images show separation and reattachment shock, as well as structures related to flow instability and transition to turbulence. A highly unsteady separation region was investigated, showing instantaneous shaping of the shock structure with turbulence.
This study explores the evolution of a turbulent hypersonic boundary layer over a spanwise-finite expansion-compression geometry. The geometry is based on a slender cone with an axial slice that subjects the cone boundary layer to a favorable pressure gradient. The mean flow field was obtained from a hybrid RANS-LES computation that showed the thickening of the boundary layer, a decrease in the mean pressure and the development of incipient streamwise vortical structures on the slice. The experiments use fluctuating surface pressure and shear-stress sensors along the centerline of the slice which demonstrate significant reduction in turbulence activity on the slice indicating relaminarization of the boundary-layer. These observations were corroborated by high framerate schlieren, filtered Rayleigh scattering and scanning focused laser differential interferometry. When a 10◦ ramp is introduced at the aft end of the slice, the effectively relaminarized boundary-layer separates upstream of the slice-ramp corner due to its increased susceptibility to separation in comparison to a turbulent boundary layer.
A new reflected shock tunnel has been commissioned at Sandia capable of generating hypersonic environments at realistic flight enthalpies. The tunnel uses an existing free-piston driver and shock tube coupled to a conical nozzle to accelerate the flow to approximately Mach 9. The facility design process is outlined and compared to other ground test facilities. A representative flight enthalpy condition is designed using an in-house state-to-state solver and piston dynamics model and evaluated using quasi-1D modeling with the University of Queensland L1d code. This condition is demonstrated using canonical models and a calibration rake. A 25 cm core flow with 4.6 MJ/kg total enthalpy is achieved over an approximately 1 millisecond test time. Analysis shows that increasing piston mass should extend test time by a factor of 2-3.
This work describes the development and testing of a carbon dioxide seeding system for the Sandia Hypersonic Wind Tunnel. The seeder injects liquid carbon dioxide into the tunnel, which evaporates in the nitrogen supply line and then condenses during the nozzle expansion into a fog of particles that scatter light via Rayleigh scattering. A planar laser scattering (PLS) experiment is conducted in the boundary layer and wake of a cone at Mach 8 to evaluate the success of the seeder. Second-mode waves and turbulence transition were well-visualized by the PLS in the boundary layer and wake. PLS in the wake also captured the expansion wave over the base and wake recompression shock. No carbon dioxide appears to survive and condense in the boundary layer or wake, meaning alternative seeding methods must be explored to extract measurements within these regions. The seeding system offers planar flow visualization opportunities and can enable quantitative velocimetry measurements in the future, including filtered Rayleigh scattering.
This experimental study explores the fluid-structure interactions occurring between a control surface and the hypersonic flow deflected by it. The control surface is simplified for this work as a spanwise finite wedge placed on a longitudinally sliced part of the cone. The front surface of the wedge is a thin panel which is designed to respond to the unsteady fluid loading arising from the shock-wave/boundary layer interactions. Experiments have been conducted in the Sandia Hypersonic Wind Tunnel at Mach 5 and Mach 8 at wedge angles of 10◦, 20◦ and 30◦ . High-speed schlieren and backside panel accelerometer measurements capture the unsteady flow dynamics and structural response of the thin panel, respectively. For attached or small separation interactions, the transitional regime has the strongest panel fluctuations with convective shock undulations induced by the boundary layer disturbance shown to be associated with dominant panel vibrations. For large separated interactions, shear layer flapping can excite select panel modes. Heating of the panel causes a downward shift in natural mode frequencies.
Measurements are presented of the aero-optic distortion produced by a Mach 8 turbulent boundary layer in the Sandia Hypersonic Wind Tunnel. Flat optical inserts installed in the test section walls enabled a double-pass arrangement of a collimated laser beam. The distortion of this beam was imaged by a high-speed Shack-Hartmann sensor at a sampling rate of up to 1 MHz. Analysis is performed using two processing methods to extract the aero-optic distortion from the data. A novel de-aliasing algorithm is proposed to extract convective-only spectra and is demonstrated to correctly quantify the physical spectra even in case of relatively low sampling rates. The results are compared with an existing theoretical model, and it is shown that this model under-predicts the experimentally measured distortions regardless of the processing method used. Possible explanations for this discrepancy are presented. The presented results represent to-date the highest Mach number for which aero-optic boundary layer distortion measurements are available.
An experimental characterization of the flow environment for the Sandia Axisymmetric Transonic Hump is presented. This is an axisymmetric model with a circular hump tested at a transonic Mach number, similar to the classic Bachalo-Johnson configuration. The flow is turbulent approaching the hump and becomes locally supersonic at the apex. This leads to a shock-wave/boundary-layer interaction, an unsteady separation bubble, and flow reattachment downstream. The characterization focuses on the quantities required to set proper boundary conditions for computational efforts described in the companion paper, including: 1) stagnation and test section pressure and temperature; 2) turbulence intensity; and 3) tunnel wall boundary layer profiles. Model characterization upstream of the hump includes: 1) surface shear stress; and 2) boundary layer profiles. Note: Numerical values characterizing the experiment have been redacted from this version of the paper. Model geometry and boundary conditions will be withheld until the official start of the Validation Challenge, at which time a revised version of this paper will become available. Data surrounding the hump are considered final results and will be withheld until completion of the Validation Challenge.
Bench-top tests are conducted to characterize Femtosecond Laser Electronic Excitation Tagging (FLEET) in static low pressure (35 mTorr-760 Torr) conditions, and to measure the acoustic disturbance caused by the resulting filament as a function of tagging wavelength and energy. The FLEET line thickness as a function of pressure and delay is described by a simple diffusion model. Initial FLEET measurements in a Mach 8 flow show that gate times of ≥ 1µs can produce visible smearing of the FLEET emission and challenge the traditional Gaussian fitting methods used to find the line center. To minimize flow perturbations and uncertainty of the final line position, several recommendations are offered: using third harmonic FLEET at 267 nm for superior signal levels with lower energy deposition than both 800 nm and 400 nm FLEET, and short camera delays and exposure times to reduce fitting uncertainty. This guidance is implemented in a Mach 8 test condition and results are presented.
Two techniques have extended the effective frequency limits of postage-stamp PIV, in which a pulse-burst laser and very small fields of view combine to achieve high repetition rates. An interpolation scheme reduced measurement noise, raising the effective frequency response of previous 400-kHz measurements from about 120 kHz to 200 kHz. The other technique increased the PIV acquisition rate to very nearly MHz rates (990 kHz) by using a faster camera. Charge leaked through the camera shift register at these framing rates but this was shown not to bias the measurements. The increased framing rate provided oversampled data and enabled use of multi-frame correlation algorithms for a lower noise floor, increasing the effective frequency response to 240 kHz where the interrogation window size begins to spatially filter the data. Good agreement between the interpolation technique and the MHz-rate PIV measurements was established. The velocity spectra suggest turbulence power-law scaling in the inertial subrange steeper than the theoretical-5/3 scaling, attributed to an absence of isotropy.
A new free-piston driven shock tube is being constructed at Sandia National Laboratories for generating extreme aerodynamic environments relevant for the study of reacting particle dispersal. The high-temperature shock tube (HST) is designed to reach post-incident shock temperatures more than 2000 K, starting from a driven section initially at ambient temperature and pressure. A design study is presented on different driver methods, leading to the selection of a free-piston driver. The tuning and performance of this driver is analyzed using the Hornung one-dimensional model and the L1d quasi-one-dimensional flow solver. The final mechanical design is shown and compared to the X2 free-piston facility. Construction was completed in mid-2018, and an initial analysis of facility performance from the first shots is presented.
Time-resolved particle image velocimetry was conducted at 40 kHz using a pulse-burst laser in the supersonic wake of a wall-mounted hemisphere. Velocity fields suggest a recirculation region with two lobes, in which flow moves away from the wall near the centerline and recirculates back toward the hemisphere off the centerline, contrary to transonic configurations. Spatio-temporal cross-correlations and conditional ensemble averages relate the characteristic behavior of the unsteady shock motion to the flapping of the shear layer. At Mach 1.5, oblique shocks develop, associated with vortical structures in the shear layer and convect downstream in tandem; a weak periodicity is observed. Shock motion at Mach 2.0 appears somewhat different, wherein multiple weak disturbances propagate from shear-layer turbulent structures to form an oblique shock that ripples as these vortices pass by. Bifurcated shock feet coalesce and break apart without evident periodicity. Power spectra show a preferred frequency of shear-layer flapping and shock motion for Mach 1.5, but at Mach 2.0, a weak preferred frequency at the same Strouhal number of 0.32 is found only for oblique shock motion and not shear-layer unsteadiness.
Femtosecond Laser Electronic Excitation Tagging (FLEET) is used to measure velocity flowfields in the wake of a sharp 7◦ half-angle cone in nitrogen at Mach 8, over freestream Reynolds numbers from 4.3∗106 /m to 13.8∗106 /m. Flow tagging reveals expected wake features such as the separation shear layer and two-dimensional velocity components. Frequency-tripled FLEET has a longer lifetime and is more energy efficient by tenfold compared to 800 nm FLEET. Additionally, FLEET lines written with 267 nm are three times longer and 25% thinner than that written with 800 nm at a 1 µs delay. Two gated detection systems are compared. While the PIMAX 3 ICCD offers variable gating and fewer imaging artifacts than a LaVision IRO coupled to a Photron SA-Z, its slow readout speed renders it ineffective for capturing hypersonic velocity fluctuations. FLEET can be detected to 25 µs following excitation within 10 mm downstream of the model base, but delays greater than 4 µs have deteriorated signal-to-noise and line fit uncertainties greater than 10%. In a hypersonic nitrogen flow, exposures of just several hundred nanoseconds are long enough to produce saturated signals and/or increase the line thickness, thereby adding to measurement uncertainty. Velocity calculated between the first two delays offer the lowest uncertainty (less than 3% of the mean velocity).
Fluid-structure interactions were studies on a 7° half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 5 and 8 and in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel. A thin composite panel was integrated into the cone and the response to boundary-layer disturbances was characterized by accelerometers on the backside of the panel. Here, under quiet-flow conditions at Mach 6, the cone boundary layer remained laminar. Artificially generated turbulent spots excited a directionally dependent panel response which would last much longer than the spot duration.
Fluid-structure interactions were studied on a 7◦ half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 5 and 8 and in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel. A thin composite panel was integrated into the cone and the response to boundary-layer disturbances was characterized by accelerometers on the backside of the panel. Under quiet-flow conditions at Mach 6, the cone boundary layer remained laminar. Artificially generated turbulent spots excited a directionally dependent panel response which would last much longer than the spot duration. When the spot generation frequency matched a structural natural frequency of the panel, resonance would occur and responses over 200 times greater than under a laminar boundary layer were obtained. At Mach 5 and 8 under noisy flow conditions, natural transition driven by the wind-tunnel acoustic noise dominated the panel response. An elevated vibrational response was observed during transition at frequencies corresponding to the distribution of turbulent spots in the transitional flow. Once turbulent flow developed, the structural response dropped because the intermittent forcing from the spots no longer drove panel vibration.