An experimental study is conducted of the fluid-thermal-structural interaction of a clamped compliant panel exposed to the intense shock-wave/boundary-layer interaction (SWBLI) induced by a compression ramp at Mach 10. Initial measurements of the underlying flowfield with a rigid ramp showed the incoming boundary layer to be transitional, and the SWBLI was observed to vary from attached to fully separated as the ramp angle was increased from 10 ^∘ to 30 ^∘ . For the compliant panel, a sealed cavity behind the panel allowed the effects of pressure-differential induced strains to be studied in the context of characterizing surface response to the aero-thermal load. Full-field, time-resolved panel deformations were measured using high-speed photogrammetry enabled by a new high-fidelity marker-tracking routine, which was shown to outperform existing methods. Substantial static panel deformations (of the order of several times the panel thickness) were produced by the intense aero-thermal loading environment. These deformations, combined with induced thermal and pressure gradients across the panel, were found to significantly modify the nature of existing panel modes (both the frequency and the displacement distributions) and introduce new, irregular mode shapes not predicted by classical clamped-plate theory; SolidWorks® simulations were performed to demonstrate that these new mode shapes were a result of the underlying panel curvature. Increasing the ramp angle resulted in a wider variety of panel modes becoming excited, while increasing the pressure differential across the panel typically produced further increases in modal frequencies and decreases in vibrational amplitudes. The transient panel response was characterized and it was found that the lower frequency mode shapes tended to gradually increase in vibrational frequency as the panel heated up and further deformed; however, higher frequency modes ( f ≳ 3 kHz ) generally showed the opposite behavior. Furthermore, as the panel deformed through the test time, the average vibrational spectra root-mean-square power was generally found to monotonically decrease.

A review is presented of recent research efforts that were conducted at the NASA Langley Research Center in the Langley Aerothermodynamics Laboratory. An overview of the test facilities, model development and fabrication capabilities, and instrumentation and measurement techniques employed in this work is provided. Contributions to hypersonic flight and planetary exploration programs are detailed, as are fundamental research and development activities. Wind tunnel investigations that supported flight programs for NASA are described herein. Collaborations between NASA projects and academia are also highlighted in this overview of recent wind tunnel experimental campaigns.

The fluid-thermal-structural interactions of a compliant panel embedded in a compression ramp are explored experimentally at Mach 6. The panel structural response is investigated for turbulent and transitional shock-wave/boundary-layer interactions at five distinct interaction strengths and up to four different pressure differentials. Upstream of the compression corner, fast-response piezoresistive pressure transducers show the feedback of the panel vibrations and static deflection on the upstream pressure fluctuations. The magnitude of the pressure fluctuation rms shows a reduction in the upstream disturbances when a pressure differential is applied across the compliant panel. Photogrammetry and IR thermography are used simultaneously to measure the panel deformations and surface temperature. Panel deformation, frequency shifting, and frequency bifurcation are consistent with post-thermal-buckling behavior.

We present results of in situ optical detection of ultrasonic waves generated in stainless steel and graphite samples mounted in a flat plate model during hypersonic flow in the 31-Inch Mach 10 wind tunnel at NASA Langley Research Center. Longitudinal waves are excited in the stainless steel and graphite sample inserts by using a contact piezoelectric transducer and the normal displacement on the surface of the sample, exposed to hypersonic fluid flow, is measured optically using a Sagnac interferometer. Measurements are consistent at different Reynolds numbers both with and without a turbulent trip strip present. Additionally, optical detection of laser-generated surface acoustic waves in a stainless-steel sample during flow is presented. These results are a demonstration of laser-based ultrasonics as an in situ material characterization technique in hypersonic flow, enabling potential applications of in situ monitoring of defect initialization and morphology in hypersonic environments.

View Video Presentation: https://doi.org/10.2514/6.2021-0601.vid Flat plate models are often used in wind tunnel testing to make measurements of the boundary layer, but preclude the use of conventional focused laser differential interferometry (FLDI) measurements, particularly for measurements at the surface of the model. By instead focusing light in only one direction using cylindrical optics, as opposed to two directions using only spherical optics, FLDI measurements can be made without truncating the beam at the model edges. Measurements at discrete points along a line can also be acquired, effectively acting as many individual FLDI instruments in a single measurement. Here, a two-line FLDI system has been constructed that allows for the measurement lines to be easily oriented at any azimuthal angle about the optical axis since the optics are housed in rotation mounts. Velocity measurements at points along a line or between two lines are also possible, since the beams are split similar to the two-point FLDI instrument. The system was first characterized in the laboratory using a laser-induced breakdown spark to provide a well-defined density fluctuation, along with high-speed schlieren measurements for comparison. The system was then installed at the NASA Langley 20-Inch Mach 6 Air Tunnel to make measurements through the boundary layer of a flat plate model and in the tunnel freestream.

Experiments were conducted to study the fluid-thermal-structural interactions of a compliant panel exposed to ramp-induced shock-wave/boundary-layer interactions (SWBLIs) at Mach 6 for laminar and transitional incoming boundary-layers. For some cases, boundary-layer trips were introduced to generate a turbulent boundary layer over half of the interrogation area. Infrared thermography was used to measure the variations in surface temperature on a compliant panel embedded in a rigid ramp. A plenum box was designed to enable the control of the back pressure behind the compliant panel. High-frequency pressure transducers measured the mean and fluctuating static pressure inside the cavity. Simultaneous temperature and pressure measurements enabled the analysis of the thermal and vibratory response of the compliant panel when subjected to SWBLIs for multiple ramp angles. The thermography also revealed an unexpected low temperature region visible on the panel at high ramp angles and Reynolds Numbers. A spanwise-nonuniform boundary layer generated by upstream trips resulted in the excitation of different modes on the panel. A study on the effects of the pressure difference on the panel natural frequencies revealed that the modes are diversely affected at low pressure differentials.

The fluid-structure interaction of a flexible panel exposed to a ramp-induced shock-wave/boundary-layer interaction (SWBLI) at Mach 6 is investigated experimentally for transitional and turbulent incoming boundary layers. Panel deformations are measured using photogrammetry enabled by a new marker-tracking routine, whereas pressure fluctuations are obtained with fast-response piezoresistive pressure transducers. The significance of aerothermal heating is evident in the nonlinear panel response: enhanced static deformations and frequency shifting are consistent with a temperature differential between the panel and its support structure, which induces compressive thermal strain and flexural softening. Time-domain and modal vibratory behavior are correlated to the SWBLI environment, and shear-layer reattachment near antinodes of certain mode shapes is identified as a source of enhanced panel excitation. Comparison with companion rigid-ramp experiments shows evidence of feedback into the downstream flowfield regime.

A multi-point focused laser differential interferometer (FLDI) has been developed to measure density fluctuations at 16 points along a line. A pair of cylindrical lenses on the transmitter side of a conventional single-point FLDI instrument form two closely spaced (≤200µm), orthogonally polarized, parallel laser lines at the instrument's focus. On the receiver side of the instrument, the interference of the beams on a 16-element photodiode array results in a single line of measurements. The further addition of a Nomarski prism creates two separate measurement lines, and the addition of a second photodiode array to the instrument enables simultaneous measurements of density fluctuations along the two lines separated by several millimeters. These two lines of measurement can be conveniently oriented at any azimuthal angle relative to the instrument's optical axis on the measurement plane, coinciding with the instrument's focus. Two experiments were performed to demonstrate the capabilities of the instrument. In the first experiment, a laser-induced breakdown spark generated a traveling spherical shock wave, and measurements of the resulting density disturbance and wave velocity were obtained. These results were compared to high-speed schlieren images of the shock wave acquired at 400 kHz. In the second experiment, the multi-point FLDI instrument was used to measure density disturbances in the boundary layer of a flat plate in a Mach 6 freestream flow. The measurements were made along two lines, both approximately 6 mm in length, extending from the surface of the plate through the boundary layer. High-speed schlieren images were acquired at 100 kHz during separate wind tunnel runs at matching unit Reynolds numbers to visualize the unsteady boundary layer flow and compare to the FLDI measurements.

A primary concern in modern fluid dynamics research is the experimental verification of computational aerothermodynamic codes. This research requires high precision and detail in the test model employed. Ceramic materials are used for these models because of their low heat conductivity and their survivability at high temperatures. To fabricate such models, slip casting techniques were developed to provide net-form, precision casting capability for high-purity ceramic materials in aqueous solutions. In previous slip casting techniques, block, or flask molds made of plaster-of-paris were used to draw liquid from the slip material. Upon setting, parts were removed from the flask mold and cured in a kiln at high temperatures. Casting detail was usually limited with this technique -- detailed parts were frequently damaged upon separation from the flask mold, as the molded parts are extremely delicate in the uncured state, and the flask mold is inflexible. Ceramic surfaces were also marred by 'parting lines' caused by mold separation. This adversely affected the aerodynamic surface quality of the model as well. (Parting lines are invariably necessary on or near the leading edges of wings, nosetips, and fins for mold separation. These areas are also critical for flow boundary layer control.) Parting agents used in the casting process also affected surface quality. These agents eventually soaked into the mold, the model, or flaked off when releasing the case model. Different materials were tried, such as oils, paraffin, and even an algae. The algae released best, but some of it remained on the model and imparted an uneven texture and discoloration on the model surface when cured. According to the present invention, a wax pattern for a shell mold is provided, and an aqueous mixture of a calcium sulfate-bonded investment material is applied as a coating to the wax pattern. The coated wax pattern is then dried, followed by curing to vaporize the wax pattern and leave a shell mold of the calcium sulfate-bonded investment material. The shell mold is cooled to room temperature, and a ceramic slip is poured therein. After a ceramic shell of desired thickness has set up in the shell mold, excess ceramic slip is poured out. While still wet, the shell mold is peeled from the ceramic shell to expose any delicate or detailed parts, after which the ceramic shell is cured to provide a complete, detailed, precision ceramic article without parting lines.

A review is presented of recent research, development, testing and evaluation activities related to entry, descent and landing that have been conducted at the NASA Langley Research Center. An overview of the test facilities, model development and fabrication capabilities, and instrumentation and measurement techniques employed in this work is provided. Contributions to hypersonic/supersonic flight and planetary exploration programs are detailed, as are fundamental research and development activities.

The analysis of effects of the reaction control system jet plumes on aftbody heating of Orion entry capsule is presented. The analysis covered hypersonic continuum part of the entry trajectory. Aerothermal environments at flight conditions were evaluated using Langley Aerothermal Upwind Relaxation Algorithm (LAURA) code and Data Parallel Line Relaxation (DPLR) algorithm code. Results show a marked augmentation of aftbody heating due to roll, yaw and aft pitch thrusters. No significant augmentation is expected due to forward pitch thrusters. Of the conditions surveyed the maximum heat rate on the aftshell is expected when firing a pair of roll thrusters at a maximum deceleration condition.

Planar laser-induced-fluorescence flowfield visualization has been used to investigate reaction-control-system jet flows in the wake of hypersonic capsule reentry vehicles. Pitch, roll, and yaw reaction-control-system jets were all studied. Planar laser-induced fluorescence was used to obtain offbody How images at planar slices in these flowfields, which are not easily visualized by other techniques, owing to characteristically low gas density. When viewed individually, these slices are shown to provide spatially and temporally resolved information, including the locations and characteristics of turbulent How structures and the location of the jet flow relative to the vehicle. In addition, ensembles of slices acquired at multiple locations throughout the flowfield are combined using computer visualization techniques to reconstruct the three-dimensional shape of the flow. Collectively, the offbody flow-visualization data set acquired in these tests represents a valuable complement to surface measurements, especially as a basis for explaining otherwise perplexing discrepancies between such measurements and computational fluid dynamics results. The tests described herein were conducted in the 31-Inch Mach 10 Air Tunnel at NASA Langley Research Center.

Pressure-sensitive and temperature-sensitive paint were used to visualize and quantify the surface interactions of reaction control system jets on the aftbody of capsule reentry vehicle shapes. The first model tested was an Apollo-like configuration and was used to focus primarily on the effects of the forward facing roll and yaw jets. The second model tested was an early Orion crew module configuration blowing only out of its forward-most yaw jet, which was expected to have the most intense aerodynamic heating augmentation on the model surface. This paper will present the results from the experiments, which show that with proper system design, both pressure-sensitive and temperature-sensitive paint are effective tools for studying these types of interaction in hypersonic testing environments.

Nitric oxide planar laser-induced fluorescence (NO PLIF) has been used to visualize the flow the aft-body of an entry capsule having an activated RCS jet in NASA Langley Research Center's 31-Inch Mach 10 wind facility. A capsule shape representative of the Apollo command module was tested. These tests were performed to demonstrate the ability of the PLIF method to visualize RCS jet flow while providing some preliminary input to NASA's Orion Vehicle design team. Two different RCS nozzle designs - conical and contoured - were tested. The conical and contoured nozzles had area ratios of 13.4 and 22.5 respectively. The conical nozzle had a half-angle of 10 . Low- and high-Reynolds number cases were investigated by changing the stagnation pressure from 350 psi to 1300 psi, resulting in freestream Reynolds numbers of 0.56 and 1.8 million per foot respectively. For both of these cases, three different jet plenum pressures were tested (nominally 56, 250 and 500 psi). A single angle-of-attack was investigated (24 degrees). NO PLIF uses an ultraviolet laser sheet to interrogate a slice in the flow containing seeded NO; this UV light excites fluorescence from the NO molecules which is detected by a high-speed digital camera. The system has spatial resolution of about 200 microns (2 pixel blurring) and has flow-stopping time resolution (approximately 1 microsecond). NO was seeded into the flow two different ways. First, the RCS jet fluid was seeded with approximately 1-5% NO, with the balance N2. This allowed observation of the shape, structure and trajectory of the RCS jets. Visualizations of both laminar and turbulent flow jet features were obtained. Visualizations were obtained with the operating at Mach 10 and also with the test section held at a constant pressure similar to the aftbody static pressure (0.04 psi) obtained during runs. These two conditions are called tunnel on and tunnel off respectively. Second, the forebody flow was seeded with a very low flowrate (<100 standard cubic centimeters per minute) of pure NO. This trace gas was entrained into and allowed visualization of the shear layer forming between the expansion fan the shoulder of the model and the recirculating separated flow in the wake of the model. This shear layer was observed to be laminar in the absence of the RCS jet operation and turbulent above a certain RCS jet flowrate. Furthermore, the operation of the RCS jet is seen to push the shear layer out away from the model, with a higher jet pressures resulting in larger deflections. Figures show some data from this test, partially processed. In the final paper, these images will be processed and rendered a three dimensional visualization of the test hardware for clearer visualization and interpretation of the flowfields.

An investigation was made in NASA Langley Research Center s 31-Inch Mach 10 Tunnel to determine the effects of reaction-control system (RCS) jet interactions on the aft-body of a capsule entry vehicle. The test focused on demonstrating and improving advanced measurement techniques that would aid in the rapid measurement and visualization of jet interaction effects for the Orion Crew Exploration Vehicle while providing data useful for developing engineering models or validation of computational tools used to assess actual flight environments. Measurements included global surface imaging with pressure and temperature sensitive paints and three-dimensional flow visualization with a scanning planar laser induced fluorescence technique. The wind tunnel model was fabricated with interchangeable parts for two different aft-body configurations. The first, an Apollo-like configuration, was used to focus primarily on the forward facing roll and yaw jet interactions which are known to have significant aft-body heating augmentation. The second, an early Orion Crew Module configuration (4-cluster jets), was tested blowing only out of the most windward yaw jet, which was expected to have the maximum heating augmentation for that configuration. Jet chamber pressures and tunnel flow conditions were chosen to approximate early Apollo wind tunnel test conditions. Maximum heating augmentation values measured for the Apollo-like configuration (>10 for forward facing roll jet and 4 for yaw jet) using temperature sensitive paint were shown to be similar to earlier experimental results (Jones and Hunt, 1965) using a phase change paint technique, but were acquired with much higher surface resolution. Heating results for the windward yaw jet on the Orion configuration had similar augmentation levels, but affected much less surface area. Numerical modeling for the Apollo-like yaw jet configuration with laminar flow and uniform jet outflow conditions showed similar heating patterns, qualitatively, but also showed significant variation with jet exit divergence angle, with as much as 25 percent variation in heat flux intensity for a 10 degree divergence angle versus parallel outflow. These results along with the fabrication methods and advanced measurement techniques developed will be used in the next phase of testing and evaluation for the updated Orion RCS configuration.

Planar laser-induced fluorescence (PLIF) has been used to investigate hypersonic flows associated with capsule reentry vehicles. These flows included reaction control system (RCS) jets, shear layer flow, and simulated forebody heatshield ablation. Pitch, roll, and yaw RCS jets were studied. PLIF obtained planar slices in these flowfields. These slices could be viewed individually or they could be combined using computer visualization techniques to reconstruct the three dimensional shape of the flow. The tests described herein were conducted in the 31-Inch Mach 10 Air Tunnel at NASA Langley Research Center. Improvements to many facets of the imaging system increased the efficiency and quality of both data acquisition, in addition to increasing the overall robustness of the system.