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Pedro Paredes

Publications and source records attributed to Pedro Paredes.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

Transition Analysis for the CRM-NLF Wind Tunnel Configuration using Transport Equation Models and Linear Stability Correlations

Transition models based on auxiliary transport equations augmenting the Reynolds-averaged Navier-Stokes (RANS) framework rely upon transition correlations that were derived from a limited number of low-speed experiments. Furthermore, these models often account for only a subset of the relevant transition mechanisms and/or cannot accurately predict the sensitivity of those mechanisms to the changes in significant flow parameters. A preceding investigation had targeted the assessment of the transport-equation-based transition models in NASA's OVERFLOW 2.3b solver, namely, the amplification factor transport (AFT-2017b) equation model coupled with the Spalart-Allmaras RANS model and the Langtry-Menter transition models (LM2009 without crossflow effects and LM2015 including the modeling of crossflow transition) implemented with Menter’s shear-stress transport equation (SST2003) RANS model. Comparisons with recent measurements at transonic freestream conditions on the Common Research Model with Natural Laminar Flow (CRM-NLF) reinforced our earlier finding that all three of the above models significantly underpredict the reported extent of the laminar flow region over the entire span of the wing, regardless of the dominant instability mechanism(s) underlying the onset of the transition process. The underprediction of the laminar flow extent was attributed to the failure of the above models in accounting for the stabilizing effect of compressibility on the amplification of Tollmien-Schlichting instabilities. Based on previous linear stability studies related to compressibility effects, the present work proposes modifications to the two classes of transition models that reduce to the original form of each model at low subsonic speeds and do not require any nonlocal flow information or additional transport equation(s). The modifications are shown to significantly improve the predicted laminar extent of the flow and compare well against the data from the CRM-NLF experiment. Additionally, a previous assessment of transition prediction based on the dual, nonparallel N -factor method in conjunction with linear parabolized stability equations (PSE) is extended to additional angles of attack to provide the first comprehensive assessment of transition models based on nonparallel disturbance amplification over the CRM-NLF. In general, the transition criterion based on the dual, nonparallel N-factor method with N TS = N CF = 6 is reasonably successful at correlating with the measured transition fronts at R eMAC = 15 million for all angles of attack investigated herein and provides additional validation of the improved predictions from the compressibility-corrected transition models.

CFD modeling↗
NASA NTRS2021

Coupling of the FUN3D Unstructured Flow Solver and the LASTRAC Stability Code to Model Transition

We develop an iterative automated method to predict transition locations in boundary-layer flows by using the FUN3D solver to perform flow simulations and the LASTRAC code for linear stability computations. The coupling of FUN3D and LASTRAC allows for a robust physics-based approach to model boundary-layer transition by analyzing the growth of different instability waves and then using that information to iteratively update the resulting transition location. There is no user involvement during the iterative computations. We apply this automated method to subsonic flow over a flat plate with a sharp leading edge. The final solution has regions of laminar and turbulent flow with a transition onset location that agrees with experiments and stability-based correlations. This iterative automated method is also applied to an NLF(1)-0416 airfoil for conditions with and without a separation bubble. Along with predicting transition locations, we compare the streamwise distributions of surface-pressure and skin-friction coefficients to a transport-equation-based model. We consider a 6:1 prolate spheroid at three angles of attack, namely, five, ten, and fifteen degrees, where mixed-mode transition occurs due to both Tollmien-Schlichting and crossflow instabilities. The skin-friction contours and transition fronts at every angle of attack from our iterative automated method show good agreement with past experimental and computational results in the literature for the 6:1 prolate spheroid.

Transition↗
NASA NTRS2021

Hypersonic Boundary-Layer Instabilities over Ogive-Cylinder Models

Computational investigations of an ogive-cylinder geometry with varying nosetips at zero degrees angle of attack are presented. The model geometry and conditions are selected to match experiments conducted in the Air Force Research Laboratory (AFRL) Mach 6 Ludwieg Tube. Five nosetips of interest were selected for the computational studies herein: two sharp ogives, two blunt ogives, and one hemispherical nosetip. Computations are performed at a freestream Reynolds number of 7.01 × 10 6 m -1 . Each sharp and blunt tip ogives had a 14 and 28 degree version for the tip angles. A cylindrical section follows the nosetip, resulting in a meter long model such as in the experiments. The laminar flow solutions are analyzed. The boundary-layer-edge properties and the velocity and temperature profiles are compared across streamwise locations aft of the ogive-cylinder junction. The blunt nosetips induce an entropy layer that envelopes the boundary-layer profiles. Modal stability analysis identifies most amplified frequencies corresponding to Mack’s second modes that agree with experimental results for the sharp tips. Similar to the experimental measurements based on wall-mounted pressure sensors, no unstable modes are found for the blunt models. Nonmodal analysis revealed a broadband set of disturbances present for the blunter tips, in agreement with experimental observations. Flow perturbation contours of most amplified planar and oblique disturbances are shown to qualitatively match wind tunnel Schlieren images, with a switch from rope-like to elongated structures, i.e., from high frequency Mack’s second modes to low frequency Mack’s first modes, as the nosetip angle is increased for the sharp tip, and from boundary-layer to entropy-layer disturbances as the bluntness is increased.

Boundary layer transition↗
NASA NTRS2021

Modeling the Effects of a Backward-Facing Step on Boundary-Layer Transition

We model transition to turbulence in a two-dimensional boundary layer downstream of a backward-facing step (BFS) along a flat plate. With the goal of evaluating the available engineering models for predicting the effects of step excrescences on the transition characteristics, two separate methodologies are used to monitor the streamwise shift in the transition onset location as the step height and the flow speed are varied across the range of a previously reported experiment involving step-height-to-local-displacement-thickness ratios of 0 < h/δ* < 1.6. Unlike the variable N -factor method from the previous literature, both of these methods are general in scope and do not involve any empirical correlations that are specific to step excrescences. The first of these techniques involves an N -factor method that directly accounts for the change in boundary-layer instability characteristics due to the step. Stability computations using the harmonic linearized Navier-Stokes equations (HLNSE), which fully account for the nonparallel-mean-flow effects close to the BFS, indicate that the measured transition locations at nearly all test conditions ( h/δ* < 1.3) correlate well with a computed N -factor of N tr = 7.6, demonstrating a successful stability-based transition criterion related to step excrescences. Linear stability theory, which does not account for nonparallel effects, demonstrates reasonable agreement with the HLNSE results, yielding good predictions for the overall trends, but predicts a somewhat earlier onset of transition than HLNSE. The other methodology used in this work involves transport-equation-based transition models. We first show that the Langtry-Menter y - Re θt transition model cannot accurately predict the location of transition onset for moderate BFS heights because it is unable to accurately account for the flow history effects. Along with the Langtry-Menter transition model, we also show the amplification factor transport model does not produce accurate transition locations for subsonic flow over steps even though it accounts for some flow history effects.

Transition↗
NASA NTRS2021

Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack

Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.

boundary layer transition↗
NASA NTRS2021

Pretest Computational Assessment of Boundary Layer Transition in the NASA Juncture Flow Model with an NACA 0015-Based Wing

The first two phases of the NASA Juncture Flow experiment were carried out on a DLR-F6 swept-wing model and were designed to provide “CFD validation-quality” data toward the assessment and improvement of existing CFD turbulence models in predicting onset and extent of three-dimensional separated flow near the wing-juncture trailing-edge region. The next phase of experiments will involve an NACA 0015-based swept wing, as prior risk reduction experiments had indicated that this wing shape resulted in reduced separation near the juncture region than the DLR-F6 wing, thus providing a better option to evaluate the ability of CFD models to predict incipient turbulent separation. The NACA 0015 measurements will also include IR thermography to infer the variation of transition front with respect to an increasing angle of attack. The primary objective of this work is to computationally make a preliminary assessment of the transition front on both surfaces of the NACA 0015 wing at a crank-chord-based Reynolds number of 2.4 x 106 for four different angles of attack, (0°, 2.5°, 5°, and 7.5°) and to determine the dominant mechanisms responsible for transition. This assessment includes both RANS-based transition models from NASA’s OVERFLOW 2.3b flow solver and linear parabolized stability equations (PSE) stability analysis based on the Langley Stability and Transition Analysis code, LASTRAC. Linear PSE results indicate that the upper surface of the wing is dominated by Tollmien- Schlichting (TS) instabilities, and that the laminar flow region shrinks from about 50% chord to a very small region just downstream of the attachment line as the angle of attack is increased from 0° to 7.5°. Consequently, the transition fronts predicted by the Spalart- Allmaras-based amplification factor transport (AFT-2017b) equation model (which accounts for the TS instabilities alone) and the Menter’s shear-stress transport equation (SST2003)- based Langtry-Menter transition model with ability to account for both TS and crossflow effects (LM2015) compare well with those predicted using linear PSE. On the lower surface of the wing, stationary crossflow (CF) instabilities begin to appear on the inboard portion of the wing in addition to the TS-instabilities for the larger angles of attack (5° and 7.5°), further reducing the laminar flow extent within the inboard region. The LM2015 model that accounts for CF effects is able to replicate this trend but appears to predict a slightly earlier transition. The outcome of this effort will inform the experiment and, when the actual experimental data become available, provide further opportunity to assess and improve the various transition models.

CFD modeling↗
NASA NTRS2021

Transition Analysis for Isolated Trips on BOLT-II Wind-Tunnel and Flight Configuration

The BOLT-II configuration designed for an upcoming flight experiment exemplifies the complexities of hypersonic boundary layer transition in the presence of multiple and potentially interacting instability mechanisms. The present paper is related to the NASA-led roughness experiment on the secondary side of the BOLT-II configuration. First, computations are performed to aid mean-flow based correlations for the transition measurements acquired during an experimental campaign in the Ludwieg Tube facility at the United States Air Force Academy. To that end, both laminar and turbulent flow computations are performed for selected conditions of the experiment with a 1:3 scale model of the slightly shorter BOLT configuration. The variation in well-known correlation parameters for transition due to isolated roughness elements with respect to element position, height, and flow Reynolds number is delineated. Furthermore, the wake instability theory is used to examine the amplification of unstable disturbances behind the centerline trip at the design conditions of the BOLT-II flight experiment and for its dynamically scaled version on the 1:3 scale model of the wind tunnel experiment. To our knowledge, this study represents the first analysis of the wake instabilities in the presence of a prominent entropy layer outside the boundary layer. At the nominal flight design condition of Re ∞ ) = 5.44 and Re ∞ = 2.5 x 10 6 /m, a diamond planform (“pizza-box”) trip with k/δ = 0.70 and planform-halfwidth-to-height ratio of b/k = 3.0 at X/L = 0.254 induces a strong distortion of the boundary-layer flow in the vicinity of the symmetry plane, such that the wake structure includes a centerline streak produced by the flow retardation behind the trip along with multiple additional, prominent vorticity structures representing the necklace vortex system. Stability results reveal that each of these streaks can support a strong amplification of wake instabilities within a frequency band of 95 kHz–255 kHz, with a peak N-factor of greater than 15 beyond X/L ≈ 0.42. The dynamically scaled wind tunnel configuration with Re ∞ = 20.25 x 10 6 /m is also found to support N > 15 at similar axial stations.

BOLT↗
NASA NTRS2021

Modeling the Effects of a Backward-Facing Step on Boundary-Layer Transition

We model transition to turbulence in a two-dimensional boundary layer downstream of a backward-facing step (BFS) along a flat plate. With the goal of evaluating the available engineering models for predicting the effects of step excrescences on the transition characteristics, two separate methodologies are used to monitor the streamwise shift in the transition onset location as the step height and the flow speed are varied across the range of a previously reported experiment involving step-height-to-local-displacement-thickness ratios of 0 < h/δ* < 1.6. Unlike the variable N -factor method from the previous literature, both of these methods are general in scope and do not involve any empirical correlations that are specific to step excrescences. The first of these techniques involves an N -factor method that directly accounts for the change in boundary-layer instability characteristics due to the step. Stability computations using the harmonic linearized Navier-Stokes equations (HLNSE), which fully account for the nonparallel-mean-flow effects close to the BFS, indicate that the measured transition locations at nearly all test conditions ( h/δ* < 1.3) correlate well with a computed N -factor of N tr = 7.6, demonstrating a successful stability-based transition criterion related to step excrescences. Linear stability theory, which does not account for nonparallel effects, demonstrates reasonable agreement with the HLNSE results, yielding good predictions for the overall trends, but predicts a somewhat earlier onset of transition than HLNSE. The other methodology used in this work involves transport-equation-based transition models. We first show that the Langtry-Menter y - Re θt transition model cannot accurately predict the location of transition onset for moderate BFS heights because it is unable to accurately account for the flow history effects. Along with the Langtry-Menter transition model, we also show the amplification factor transport model does not produce accurate transition locations for subsonic flow over steps even though it accounts for some flow history effects.

Transition↗
NASA NTRS2021

Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack

Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.

Boundary Layer Transition↗
NASA NTRS2021

Pretest Computational Assessment of Boundary Layer Transition in the NASA Juncture Flow Model with an NACA 0015-Based Wing

The first two phases of the NASA Juncture Flow experiment were carried out on a DLR-F6 swept-wing model and were designed to provide “CFD validation-quality” data toward the assessment and improvement of existing CFD turbulence models in predicting onset and extent of three-dimensional separated flow near the wing-juncture trailing-edge region. The next phase of experiments will involve an NACA 0015-based swept wing, as prior risk reduction experiments had indicated that this wing shape resulted in reduced separation near the juncture region than the DLR-F6 wing, thus providing a better option to evaluate the ability of CFD models to predict incipient turbulent separation. The NACA 0015 measurements will also include IR thermography to infer the variation of transition front with respect to an increasing angle of attack. The primary objective of this work is to computationally make a preliminary assessment of the transition front on both surfaces of the NACA 0015 wing at a crank-chord-based Reynolds number of 2.4 x 106 for four different angles of attack, (0°, 2.5°, 5°, and 7.5°) and to determine the dominant mechanisms responsible for transition. This assessment includes both RANS-based transition models from NASA’s OVERFLOW 2.3b flow solver and linear parabolized stability equations (PSE) stability analysis based on the Langley Stability and Transition Analysis code, LASTRAC. Linear PSE results indicate that the upper surface of the wing is dominated by Tollmien- Schlichting (TS) instabilities, and that the laminar flow region shrinks from about 50% chord to a very small region just downstream of the attachment line as the angle of attack is increased from 0° to 7.5°. Consequently, the transition fronts predicted by the Spalart- Allmaras-based amplification factor transport (AFT-2017b) equation model (which accounts for the TS instabilities alone) and the Menter’s shear-stress transport equation (SST2003)- based Langtry-Menter transition model with ability to account for both TS and crossflow effects (LM2015) compare well with those predicted using linear PSE. On the lower surface of the wing, stationary crossflow (CF) instabilities begin to appear on the inboard portion of the wing in addition to the TS-instabilities for the larger angles of attack (5° and 7.5°), further reducing the laminar flow extent within the inboard region. The LM2015 model that accounts for CF effects is able to replicate this trend but appears to predict a slightly earlier transition. The outcome of this effort will inform the experiment and, when the actual experimental data become available, provide further opportunity to assess and improve the various transition models.

CFD modeling↗
NASA NTRS2021

Transition Analysis for Isolated Trips on BOLT-II Wind-Tunnel and Flight Configuration

The BOLT-II configuration designed for an upcoming flight experiment exemplifies the complexities of hypersonic boundary layer transition in the presence of multiple and potentially interacting instability mechanisms. The present paper is related to the NASA-led roughness experiment on the secondary side of the BOLT-II configuration. First, computations are performed to aid mean-flow based correlations for the transition measurements acquired during an experimental campaign in the Ludwieg Tube facility at the United States Air Force Academy. To that end, both laminar and turbulent flow computations are performed for selected conditions of the experiment with a 1:3 scale model of the slightly shorter BOLT configuration. The variation in well-known correlation parameters for transition due to isolated roughness elements with respect to element position, height, and flow Reynolds number is delineated. Furthermore, the wake instability theory is used to examine the amplification of unstable disturbances behind the centerline trip at the design conditions of the BOLT-II flight experiment and for its dynamically scaled version on the 1:3 scale model of the wind tunnel experiment. To our knowledge, this study represents the first analysis of the wake instabilities in the presence of a prominent entropy layer outside the boundary layer. At the nominal flight design condition of Re ∞ ) = 5.44 and Re ∞ = 2.5 x 10 6 /m, a diamond planform (“pizza-box”) trip with k/δ = 0.70 and planform-halfwidth-to-height ratio of b/k = 3.0 at X/L = 0.254 induces a strong distortion of the boundary-layer flow in the vicinity of the symmetry plane, such that the wake structure includes a centerline streak produced by the flow retardation behind the trip along with multiple additional, prominent vorticity structures representing the necklace vortex system. Stability results reveal that each of these streaks can support a strong amplification of wake instabilities within a frequency band of 95 kHz–255 kHz, with a peak N-factor of greater than 15 beyond X/L ≈ 0.42. The dynamically scaled wind tunnel configuration with Re ∞ = 20.25 x 10 6 /m is also found to support N > 15 at similar axial stations.

BOLT↗
NASA NTRS2021

Assessment of RANS-based Transition Models based on Experimental Data of the Common Research Model with Natural Laminar Flow

Transition models based on auxiliary transport equations augmenting the Reynolds-averaged Navier-Stokes (RANS) framework often rely upon transition correlations that were derived from a limited number of low-speed experiments and these models often fail to account for all of the relevant transition mechanisms and/or the variation in those mechanisms with respect to changes in the significant flow parameters. Available data from a recent experiment on the Common Research Model with Natural Laminar Flow (CRM-NLF) in the National Transonic Facility at the NASA Langley Research Center are used to assess the current transition modeling capability in NASA's OVERFLOW 2.3b solver for a swept wing configuration with nonzero taper and transonic cruise conditions. Specifically, the OVERFLOW solutions are used to evaluate the accuracy and robustness of the transport-equation-based transition models. Results highlight that the Spalart-Allmaras-based amplification factor transport (AFT-2017b) equation model and Menter’s shear-stress transport equation (SST2003)-based Langtry-Menter transition models (either with or without the modeling of crossflow transition) significantly underpredict the reported extent of laminar flow region over the entire span of the wing, irrespective of which instability mechanism(s) is expected to dominate the onset of the transition process. We show that the transition correlations underlying these models fail to account for the stabilizing effect of compressibility on the Tollmien-Schlichting transition, which is likely to be a major contributor to the underprediction of the laminar flow region on the CRM-NLF. The SST-2003-based Langtry-Menter model also appears to inaccurately predict the chordwise pressure variation along the majority of the wing span at all the flow conditions studied herein, due to how the turbulence intensity levels were enforced in the computations and how that was interfering with the functioning of the underlying turbulence model within the boundary layer. The AFT and Langtry-Menter models appear to be sensitive to the level of the freestream turbulence intensity, but the degree of sensitivity varies across the models.

CFD modeling↗
NASA NTRS2021

A Computational Analysis of Boundary Layer Instability over the BOLT Configuration (Effect of Nonzero AoA and Yaw)

The complex boundary layer flow over the BOLT flight configuration is known to exhibit multiple and potentially interacting instability mechanisms. This paper represents a continuation of our numerical investigation of the flow instabilities over the main test surface of the BOLT configuration by using state-of-the-art tools in multidimensional instability analysis. Specifically, the paper extends our previous computations by considering the separate effects of a nonzero angle of attack and a nonzero yaw on the modal instability characteristics of the boundary layer streaks adjacent to the minor-axis symmetry plane, specifically near both ends of the azimuthal region of a thick boundary layer in the middle, where this region rapidly changes to a thinner boundary layer on either side. At the t = 28.8767 s condition from the ascent part of the anticipated flight trajectory with a flight Mach number of M ∞ = 5.53 and unit Reynolds number of 4.25 x 10 6 /m, either type of departure from the design condition is shown to have a considerable impact on the structure of the basic state rollup within the region of interest and, hence, also on the amplification characteristics of instability waves within the resulting streaks. Yet, for a yaw angle of β = 4 degrees, the computations indicate only a slight reduction with respect to the peak N-factor of nearly 11 at the design condition of zero degrees yaw and zero degrees angle of attack. In contrast, an angle of attack equal to α = 4 degrees, the peak N-factor decreases to nearly 6 on the leeward side and increases above 16 on the windward side, making the onset of transition highly likely on the windward side. Computations also highlight the role of streak instabilities that originate as Mack mode disturbances and also demonstrate the potential pitfalls in using surface pressure sensors alone to gauge the magnitude of instability amplification.

Boundary layer transition↗
NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗
NASA NTRS2021

Hypersonic Second Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗