Gal4 turnover and transcription activation

Growing evidence supports the notion that proteasome-mediated destruction of transcriptional activators can be intimately coupled to their function. Recently, Nalley et al. challenged this view by reporting that the prototypical yeast activator Gal4 does not dynamically associate with chromatin, but rather 'locks in' to stable promoter complexes that are resistant to competition. Here we present evidence that the assay used to reach this conclusion is unsuitable, and that promoter-bound, active Gal4 is indeed susceptible to competition in vivo. Our data challenge the key evidence that Nalley et al. used to reach their conclusion, and indicate that Gal4 functions in vivo within the context of dynamic promoter complexes

Using Pressure- and Temperature-Sensitive Paint for Global Surface Pressure and Temperature Measurements on the Aft-Body of a Capsule Reentry Vehicle

Pressure Sensitive Paint (PSP) and Temperature Sensitive Paint (TSP) were used to visualize and quantify the surface interactions of reaction control system (RCS) jets on the aft body 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 PSP and TSP are effective tools for studying these types of interaction in hypersonic testing environments

Shockwave/Boundary-Layer Interaction Studies Performed in the NASA Langley 20-Inch Mach 6 Air Tunnel

This paper highlights results from a collaborative study performed by The University of Tennessee Space Institute (UTSI) and NASA Langley Research Center on the Shockwave/Boundary-Layer Interaction (SWBLI) generated by a cylindrical protuberance on a flat plate in a Mach 6 flow. The study was performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center and consisted of two separate entries. In the first entry, simultaneous high-speed schlieren and high-speed pressure-sensitive paint (PSP) imaging which was performed for the first time in the 20-Inch Mach 6 facility at NASA Langley were performed as well as simultaneous high-speed schlieren and oil-flow imaging. In the second entry, the model configuration was modified to increase the size of the interaction region. High-speed schlieren and infrared thermography (IR) surface imaging were performed in this second entry. The goal of these tests was to characterize the SBLI in the presence of a laminar, transitional, and turbulent boundary layer using high-speed optical imaging techniques. AoA = sting angle-of-attack () dcylinder = cylinder diameter (mm) dtrip = cylindrical tripping element diameter (mm) shock = shock stand-off distance (mm) hcylinder = cylinder height (mm) htrip = cylindrical tripping element height (mm) HSS = high-speed schlieren M = freestream Mach number PSP = pressure-sensitive paint Re = freestream unit Reynolds number (m-1) SWBLI = shockwave/boundary-layer interaction plate = model plate angle () Introduction his paper highlights two experimental entries performed in the 20-Inch Mach 6 Air Blowdown Tunnel at NASA Langley Research Center in collaboration with The University of Tennessee Space Institute (UTSI). The purpose of these entries was to characterize the dynamic shockwave/boundary-layer interaction (SWBLI) between a vertical cylinder on a flat plate and laminar, transitional (XSWBLI), and turbulent (SWTBLI) boundary layers with a freestream Mach number of 6 using non-intrusive optical diagnostics. Experiments performed by Murphree et al.1,2 were among the first to specifically characterize XSWBLI induced by a vertical cylinder on a flat plate geometry using several optical measurement techniques. Recent optical studies of XSWBLI phenomenon have been performed by UTSI at Mach 2 in their low-enthalpy blow wind tunnel3-8 and by Texas A&M University and UTSI at Mach numbers of 6 and 7 in their Adjustable Contour Expansion wind tunnel.9 The experiments described in this paper were intended to complement previous studies by expanding the freestream unit Reynolds number range, Re, over which the XSWBLI phenomena has been observed. Additionally these experiments, made possible under NASAs new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project, promoted collaboration between university and NASA researchers. The initial entry in the 20-Inch Mach 6 Air Tunnel at NASA Langley occurred in December of 2016. Originally, testing was to occur in November of 2016 in the 31-Inch Mach 10 Air Tunnel at NASA Langley. This facility was chosen so that the XSWBLI phenomenon could be observed at much higher Mach numbers than had previously been attempted in ground test experiments. The model selected for this experiment, a 10 half-angle wedge with a sharp leading edge (described in detail in section II.B), had previously been used by Danehy et al. [10] for boundary layer transition studies using the nitric oxide planar laser-induced fluorescence (NO PLIF) flow visualization technique. In that work, it was determined that transition could be induced downstream of a single htrip = 1-mm tall, dtrip = 4-mm diameter cylindrical tripping element and that the streamwise location of the transition could be changed for a single Re by changing the model angle-of-attack (AoA) (see Fig. A3 in Ref. [10] for more details). Based on the findings of that work, a decision was made to use the wedge model with the cylindrical tripping element to trip the boundary layer flow ahead of a cylindrical protuberance in order to achieve a XSWBLI. Unfortunately, the 31-Inch Mach 10 facility had been taken offline for repairs in October of 2016 and a decision was made to move the test to the 20-Inch Mach 6 facility. Since the behavior of the boundary layer with the chosen model configuration had not been studied before in that facility and the available test time was limited, the entry was considered to be exploratory and was used to collect spatially-resolved and time-resolved flow and surface visualization data that would be used to inform a second entry. Test techniques included simultaneous high-speed schlieren (HSS) captured at 160 kHz and high-speed pressure sensitive paint captured at 10 kHz as well as oil flow visualization, captured at 750 Hz. The second entry in the 20-Inch Mach 6 facility occurred in June and July of 2017. In this follow-on test, modifications to the wind tunnel model were made based on observations made during the first entry and included removing the cylindrical tripping element, increasing the size of the cylinder used to induce the SWBLI to increase the size of the interaction while simultaneously improving spatial resolution, and using a swept ramp array, similar to that described in Ref. [11], to trip the flow to turbulence. Simultaneous HSS (captured at 140 kHz, 100 kHz, and 40 kHz) and conventional IR thermography (captured at 30 Hz) imaging were performed simultaneously in this follow-on entry. This paper is intended to serve as a summary of the work performed during these two entries, to detail lessons learned from each entry, and to highlight some of the datasets acquired. Details on the experimental setup, model configuration, and techniques used are provided. Papers providing a more rigorous analysis of data acquired during the second entry, including statistical, spectral, and modal decomposition methods, can be found in Refs. [12,13]. An entry examining XSWBLI in the 31-Inch Mach 10 Blowdown Wind Tunnel facility is currently planned for mid-to-late calendar year 2019, pending the success of facility repairs. The work performed and described in this paper and the upcoming entry in the 31-Inch Mach 10 facility at NASA Langley have been made possible by NASAs new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project. Wind Tunnel Facility All experiments discussed in this paper were performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center. Specific details pertaining to this facility can be found in Refs. [14,15], with only a brief description of the facility provided here. For both entries, the nominal freestream unit Reynolds number was varied between 1.8106 m-1 (0.5106 ft-1) and 26.3106 m-1 (8106 ft-1). The nominal stagnation pressure was varied between 0.21 MPa and 3.33 MPa and the nominal stagnation temperature was varied between 480 K and 520 K to achieve the desired Re condition. For all runs, the nominal freestream Mach number was 6. The nearly square test section is 520.7-mm (20.5-inches) wide by 508-mm (20-inches) high. Two 431.8-mm (17-inch) diameter windows made of Corning 7940, Grade 5F schlieren-quality glass serve as the side walls of the tunnel and provide optical access for the high-speed schlieren measurements. A rectangular window made of the same material as the side windows served as the top wall of the test section and provided optical access for the high-speed PSP and oil flow measurements. For the second entry, this top window was replaced with a Zinc Selenide (ZnSe) window with an anti-reflection coating capable of passing IR wavelengths between 8m and 12m with greater than 98% transmittance. The model was sting supported by a strut attached to a hydraulic system that allows for the model pitch angle to be adjusted between -5 to +55. For the first entry, an initial pitch/pause sweep of the model AoA was performed to observe the resulting SWBLI. Ultimately, however, the sting pitch angle for this entry was fixed at +10.0 so that the angle of the top surface of the wedge relative to the streamwise axis of the tunnel (referred to herein as the plate angle, plate), was plate = 0. For the second entry, plate = 0 and plate = -13.25 were initially tested with the swept ramp array (discussed in the following section) to determine which orientation produced conditions most favorable for XSWBLI to occur based on the heating signatures observed over the top surface of the model in the IR thermography images. Based on these initial tests, plate = -13.25 was set for the remainder of the runs in the second entry. For both entries, any model changes were performed in a housing located beneath the closed test section. Prior to performing a run of the tunnel, the housing was sealed and the tunnel started. Once the appropriate freestream conditions were achieved, the model was injected into the test section using a hydraulic injection system. B. Model Geometry For all runs, a 10 half-angle (20 full-angle) wedge model with a sharp leading edge was used. The model is described in detail in Refs. [10,16]. The top surface of the sharp leading edge of the model extended 47.8 mm from its upstream-most edge to a junction with the upstream edge of a stainless steel top plate that then extended an (a) (c) (b) Fig. 1 (a) Schematic of top surface of wedge model with gas seeding insert, (b) perspective view of the model in the 20-Inch Mach 6 tunnel with centerline pressure orifices on sharp leading edge, and (c) a perspective view of the model with stainless steel (top) and SLA middle insert (bottom) during the first entry. Flow occurs from left to right

Using Pressure Sensitive Paint to Measure Aerodynamic Forces on a Rotor Blade in Hover

This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on rotorcraft blades in hover at the Rotor Test Cell located in the 14- by 22-Foot Subsonic Tunnel complex at the NASA Langley Research Center. This work builds upon previous entries and focused on collecting measurements from the upper and lower surface simultaneously. From these results, normal force (F (sub z)) values can be obtained. To date, this is the first time that the Pressure Sensitive Paint technique has been used for these types of measurements on rotor blades. In addition, several areas of improvement have been identified and are currently being developed for future testing

Development of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades

This paper will describe the results from a proof of concept test to examine the feasibility of using Pressure Sensitive Paint (PSP) to measure global surface pressures on rotorcraft blades in hover. The test was performed using the U.S. Army 2-meter Rotor Test Stand (2MRTS) and 15% scale swept rotor blades. Data were collected from five blades using both the intensity- and lifetime-based approaches. This paper will also outline several modifications and improvements that are underway to develop a system capable of measuring pressure distributions on up to four blades simultaneously at hover and forward flight conditions

Head-Neck Dual-energy CT Contrast Media Reduction Using Diffusion Models

Iodinated contrast media is essential for dual-energy computed tomography (DECT) angiography. Previous studies show that iodinated contrast media may cause side effects, and the interruption of the supply chain in 2022 led to a severe contrast media shortage in the US. Both factors justify the necessity of contrast media reduction in relevant clinical applications. In this study, we propose a diffusion model-based deep learning framework to address this challenge. First, we simulate different levels of low contrast dosage DECT scans from the standard normal contrast dosage DECT scans using material decomposition. Conditional denoising diffusion probabilistic models are then trained to enhance the contrast media and create contrast-enhanced images. Our results demonstrate that the proposed methods can generate high-quality contrast-enhanced results even for images obtained with as low as 12.5% of the normal contrast dosage. Furthermore, our method outperforms selected competing methods in a human reader study

Applying Pressure Sensitive Paint Technology to Rotor Blades

This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on rotorcrtaft blades in simulated forward flight at the 14- by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The basics of the PSP method will be discussed and the modifications that were needed to extend this technology for use on rotor blades. Results from a series of tests will also be presented as well as several areas of improvement that have been identified and are currently being developed for future testing

Hypersonic Laminar Boundary Layer Velocimetry with Discrete Roughness on a Flat Plate

Laminar boundary layer velocity measurements are made on a 10-degree half-angle wedge in a Mach 10 flow. Two types of discrete boundary layer trips were used to perturb the boundary layer gas. The first was a 2-mm tall, 4-mm diameter cylindrical trip. The second was a scaled version of the Orbiter Boundary Layer Transition (BLT) Detailed Test Objective (DTO) trip. Both 1-mm and 2.5-mm tall BLT DTO trips were tested. Additionally, side-view and plan-view axial boundary layer velocity measurements were made in the absence of these tripping devices. The free-stream unit Reynolds numbers tested for the cylindrical trips were 1.7x10(exp 6)/m and 3.3x10(exp 6)/m. The free-stream unit Reynolds number tested for the BLT DTO trips was 1.7x10(exp 6)/m. The angle of attack was kept at approximately 5-degrees for most of the tests resulting in a Mach number of approximately 8.3. These combinations of unit Reynolds numbers and angle of attack resulted in laminar flowfields. To study the precision of the measurement technique, the angle of attack was varied during one run. Nitric-oxide (NO) molecular tagging velocimetry (MTV) was used to obtain averaged axial velocity values and associated uncertainties. These uncertainties are as low as 20 m/s. An interline, progressive scan CCD camera was used to obtain separate images of the initial reference and shifted NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond sequential acquisition of both images. The maximum planar spatial resolution achieved for the side-view velocity measurements was 0.07-mm in the wall-normal direction by 1.45-mm in the streamwise direction with a spatial depth of 0.5-mm. For the plan-view measurements, the maximum planar spatial resolution in the spanwise and streamwise directions was 0.69-mm by 1.28-mm, respectively, with a spatial depth of 0.5-mm. Temperature sensitive paint (TSP) measurements are provided to compliment the velocity data and to provide further insight into the behavior of the boundary layers. The experiments were performed at the NASA Langley Research Center 31-Inch Mach 10 Air tunnel

Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight

This paper will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. Preliminary results show that the PSP agrees both qualitatively and quantitatively with both the expected results as well as with the pressure taps. Several areas of improvement have been indentified and are currently being developed

Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight: Preliminary PSP Results from Test 581 in the 14- by 22-Foot Subsonic Tunnel

This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. This report will also detail possible improvements to the system