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Judith Hannon

Publications and source records attributed to Judith Hannon.

10 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

A Simple Boundary-Layer Transition Detection Method Using Mean Static Pressure Measurements

A simple method for detecting boundary-layer transition using only mean static pressure port data is presented. The method can be applied to most existing models with pressure taps, and only requires that a fine angle-of-attack sweep be performed. A small but abrupt change in the static pressure is visible when the transition front passes over the pressure tap. Results from a recent Juncture Flow test entry are used to illustrate the technique. Infrared thermography measurements of the transition front compare very well to the transition locations obtained from the static pressure ports. Some differences in behavior occur depending on the dominant transition mechanism. While the technique is somewhat qualitative, it can be an excellent tool for estimating the transition location when other tools are not readily available. Given the simplicity of the technique, and the fact that most wind-tunnel models are already designed with numerous static pressure taps, this method can be applied with very little overhead.

boundary-layer transition↗
NASA NTRS2023

Measurements and Computations of Natural Transition on the NASA Juncture-Flow Model with a Symmetric Wing

Experiments were performed in the 14- by 22-Foot Subsonic Tunnel to assess natural transition on the symmetric-airfoil wings of the NASA Juncture-Flow Model. Infrared thermography was used to visualize the heating on the upper surface of both wings of the full-span model, and on the fuselage, for angles of incidence ranging from -10° to 10° at a fixed Reynolds number of 2.4E6 based on the chord length at the wing planform break. The fuselage boundary layer transitioned well upstream of the wing-root leading edge for all conditions. Transition fronts were identified by a steep rise in the surface temperature, and the transition coordinates were transformed from an image-based to a body-fixed system. Additionally, the state of the boundary layer was estimated at pressure ports distributed on the wings through observation of the pressure coefficient as a function of the angle of incidence. For increasing angles of incidence, the transition front was observed to advance upstream, in a mostly spanwise-uniform fashion, from near midchord at α = 0°; however, for increasingly negative angles of incidence, the transition front first receded and then advanced in a nonuniform jagged manner that is typically observed with stationary crossflow. The transition wedges first appeared inboard of the wing break and then spread outboard to near the tip by α = -6°. The upstream shift in transition at positive angles of incidence and the outboard progression of crossflow-dominated transition at increasingly negative angles of incidence are consistent with trends identified in a computational assessment of the boundary-layer transition based on both linear stability analysis and Reynolds-averaged-Navier-Stokes-based transition models. The stability results obtained from the Langley Stability and Transition Analysis Code were used to recalibrate a dual N-factor criterion, which allowed for the prediction of transition fronts that showed excellent agreement with the experiment. The Reynolds-averaged-Navier-Stokes-based models, from the NASA OVERFLOW 2.3 solver, that accounted for the crossflow instability showed mixed results in comparison with the experiment, with the helicity-based Langtry-Menter model performing the best. The experimental data, particularly the cases involving strong influence from both Tollmien-Schlichting and crossflow instabilities, will be valuable for the continued validation and improvement of transition models.

boundary layer transition↗
NASA NTRS2023

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗
NASA NTRS2023

Measurements and Computations of Natural Transition on the NASA Juncture-Flow Model with a Symmetric Wing

Experiments were performed in the 14- by 22-Foot Subsonic Tunnel to assess natural transition on the symmetric-airfoil wings of the NASA Juncture-Flow Model. Infrared thermography was used to visualize the heating on the upper surface of both wings of the full-span model, and on the fuselage, for angles of incidence ranging from -10° to 10° at a fixed Reynolds number of 2.4E6 based on the chord length at the wing planform break. The fuselage boundary layer transitioned well upstream of the wing-root leading edge for all conditions. Transition fronts were identified by a steep rise in the surface temperature, and the transition coordinates were transformed from an image-based to a body-fixed system. Additionally, the state of the boundary layer was estimated at pressure ports distributed on the wings through observation of the pressure coefficient as a function of the angle of incidence. For increasing angles of incidence, the transition front was observed to advance upstream, in a mostly spanwise-uniform fashion, from near midchord at α = 0°; however, for increasingly negative angles of incidence, the transition front first receded and then advanced in a nonuniform jagged manner that is typically observed with stationary crossflow. The transition wedges first appeared inboard of the wing break and then spread outboard to near the tip by α = -6°. The upstream shift in transition at positive angles of incidence and the outboard progression of crossflow-dominated transition at increasingly negative angles of incidence are consistent with trends identified in a computational assessment of the boundary-layer transition based on both linear stability analysis and Reynolds-averaged-Navier-Stokes-based transition models. The stability results obtained from the Langley Stability and Transition Analysis Code were used to recalibrate a dual N-factor criterion, which allowed for the prediction of transition fronts that showed excellent agreement with the experiment. The Reynolds-averaged-Navier-Stokes-based models, from the NASA OVERFLOW 2.3 solver, that accounted for the crossflow instability showed mixed results in comparison with the experiment, with the helicity-based Langtry-Menter model performing the best. The experimental data, particularly the cases involving strong influence from both Tollmien-Schlichting and crossflow instabilities, will be valuable for the continued validation and improvement of transition models.

Boundary layer transition↗
NASA NTRS2023

Measurements and Computations of Natural Transition on the NASA Juncture-Flow Model with a Symmetric Wing

Experiments were performed in the 14- by 22-Foot Subsonic Tunnel to assess natural transition on the symmetric-airfoil wings of the NASA Juncture-Flow Model. Infrared thermography was used to visualize the heating on the upper surface of both wings of the full-span model, and on the fuselage, for angles of incidence ranging from -10° to 10° at a fixed Reynolds number of 2.4E6 based on the chord length at the wing planform break. The fuselage boundary layer transitioned well upstream of the wing-root leading edge for all conditions. Transition fronts were identified by a steep rise in the surface temperature, and the transition coordinates were transformed from an image-based to a body-fixed system. Additionally, the state of the boundary layer was estimated at pressure ports distributed on the wings through observation of the pressure coefficient as a function of the angle of incidence. For increasing angles of incidence, the transition front was observed to advance upstream, in a mostly spanwise-uniform fashion, from near midchord at α = 0°; however, for increasingly negative angles of incidence, the transition front first receded and then advanced in a nonuniform jagged manner that is typically observed with stationary crossflow. The transition wedges first appeared inboard of the wing break and then spread outboard to near the tip by α = -6°. The upstream shift in transition at positive angles of incidence and the outboard progression of crossflow-dominated transition at increasingly negative angles of incidence are consistent with trends identified in a computational assessment of the boundary-layer transition based on both linear stability analysis and Reynolds-averaged-Navier-Stokes-based transition models. The stability results obtained from the Langley Stability and Transition Analysis Code were used to recalibrate a dual N-factor criterion, which allowed for the prediction of transition fronts that showed excellent agreement with the experiment. The Reynolds-averaged-Navier-Stokes-based models, from the NASA OVERFLOW 2.3 solver, that accounted for the crossflow instability showed mixed results in comparison with the experiment, with the helicity-based Langtry-Menter model performing the best. The experimental data, particularly the cases involving strong influence from both Tollmien-Schlichting and crossflow instabilities, will be valuable for the continued validation and improvement of transition models.

boundary layer transition↗
NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗
NASA NTRSDate not supplied

Active Flow Control Enhanced Aileron of the High-Lift Common Research Model at Takeoff Condition

An experimental investigation was conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel on the NASA 10% scale High-Lift Common Research Model (CRM-HL) to improve the high-lift aerodynamic performance of the takeoff configuration. Active flow control (AFC) was applied at the aileron hinge line to control flow separation at high aileron deflection angles ( δ a ) to increase lift-to-drag ratio (R/J). Improvements in low speed R/J can increase vehicle payload and/or range resulting in a more efficient vehicle. Aileron deflection angles of 0°, 7.5°, 16°, and 25° were investigated. Flow control was applied to the two largest aileron deflections, which are considered the high-lift (HL) improved takeoff configurations. A companion paper will discuss baseline flow characteristics of the CRM-HL at the three takeoff conditions: reference, nominal, and HL-improved. In this paper, we focus on the aerodynamic improvement obtained when AFC was applied to δ a = 16° relative to the nominal configuration, δ a = 7.5°. Multiple flow control parameters, including actuator type, spacing, and intensity, were investigated to evaluate the efficiency of the actuation system. The different AFC configurations tested were assessed with tuft visualization, steady surface pressure data, and force and moment data. The results indicated that all actuation types and spacings examined led to an increase in lift across the range of angles of attack investigated. Flow separation was mitigated with the injection of momentum at the aileron hinge line. The configurations with the smallest spacing produced the highest improvement in L/D . Steady jet actuation showed higher efficiency but similar aerodynamic performance when compared to the sweeping jet actuators. This work was performed in support of the NASA Advanced Air Transport Technology (AATT) Project.

High-Lift Common Research Model↗
NASA NTRSDate not supplied

Testing of High-Lift Common Research Model at Takeoff Configurations

The 10% scale High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The main objective of the wind tunnel test was to improve the aerodynamic performance of a representative aircraft model during takeoff operations using localized active flow control (AFC). This approach involves the application of AFC concepts locally to a relatively small region; therefore, it has the potential to cause minimal architectural change to current aircraft configurations. In addition, the power requirements of localized AFC could be supplied with onboard air resources. Recent exploratory studies identified the aileron as a target area for improving lift-to-drag ratio using localized AFC concepts. The idea is to deflect the ailerons beyond their nominal deflection angles and use AFC to reduce flow separation that occurs at larger aileron deflections. Several AFC concepts with different configurations were evaluated with the goal of achieving high-lift performance improvement. The assessment of AFC configurations is reported in a companion paper. The focus of the current paper is to report the data relevant to the CRM-HL takeoff configurations and establish a reference case for localized AFC application. Three takeoff configurations —reference, nominal, and high-lift improved—are documented. These takeoff configurations are obtained by varying aileron deflection. Wind tunnel measurements including surface static pressures and force and moment data are presented. In addition, surface tuft flow visualization, mainly on the aileron, is provided to understand the flow characteristics developed over the aileron during takeoff.

Wind tunnel test↗
NASA NTRSDate not supplied

Active Flow Control Enhanced Aileron of the High-Lift Common Research Model at Takeoff Condition

An experimental investigation was conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel on the NASA 10% scale High-Lift Common Research Model (CRM-HL) to improve the high-lift aerodynamic performance of the takeoff configuration. Active flow control (AFC) was applied at the aileron hinge line to control flow separation at high aileron deflection angles to increase lift-to-drag ratio (L/D). Improvements in low speed L/D can increase vehicle payload and/or range resulting in a more efficient vehicle. Aileron deflection angles of 0°, 7.5°, 16°, and 25° were investigated. Flow control was applied to the two largest aileron deflections, which are considered the high-lift (HL) improved takeoff configurations. A companion paper will discuss baseline flow characteristics of the CRM-HL at the three takeoff conditions: reference, nominal, and HL-improved. In this paper, we focus on the aerodynamic improvement obtained when AFC was applied to an aileron deflected at 16° relative to the nominal configuration of 7.5° deflection. Multiple flow control parameters, including actuator type, spacing, and intensity, were investigated to evaluate the efficiency of the actuation system. The different AFC configurations tested were assessed with tuft visualization, steady surface pressure data, and force and moment data. The results indicated that all actuation types and spacings examined led to an increase in lift across the range of angles of attack investigated. Flow separation was mitigated with the injection of momentum at the aileron hinge line. The configurations with the smallest spacing produced the highest improvement in L/D. Steady jet actuation showed higher efficiency but similar aerodynamic performance when compared to the sweeping jet actuators. This work was performed in support of the NASA Advanced Air Transport Technology (AATT) Project.

CRM-HL↗