Search NASA⌕ Search

SEARCH · Search NASA

Results for “Unitary Plan Wind Tunnel”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

392 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Camera Setup for Unsteady Pressure Sensitive Paint at NASA Ames Research Center

The unsteady Pressure-Sensitive Paint Development Team at NASA Ames Research Center is conducting a Launch Vehicle Demo Test in the Unitary Plan Wind Tunnel’s 11-ft Transonic Wind Tunnel in the first half of 2024 to showcase improvements made to unsteady Pressure-Sensitive Paint technology. The unsteady Pressure-Sensitive Paint setup has upgraded from a four-camera system to an eight-camera system, prompting analysis to maximize the system performance and overall data quality. Additional analysis was performed to maximize the system resolving power when subject to test constraints such as model size and model tunnel positions. These analyses have uncovered new opportunities to improve the unsteady Pressure-Sensitive Paint technology and will allow the system to collect higher quality data with simple changes to the setup procedure.

Camera Calibration↗
NASA NTRS2024

Comparison of OVERFLOW Computational and Experimental Results of the CobraMRV Mars Entry Vehicle Concept during Supersonic Retropropulsion

The CFD solver OVERFLOW was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a CFD and wind-tunnel integration effort. The aerodynamics of the CobraMRV are summarized and comparisons to experimental data at flight conditions of interest are presented. The sensitivity of vehicle loads and flow field unsteadiness to CFD parameters at a subset of operating conditions are investigated.

CFD↗
NASA NTRS2024

OVERFLOW Guidelines for Simulation of Supersonic Retropropulsion

General methods for obtaining satisfactory simulation accuracy across a range of supersonic retropropulsive flows are recommended using comparisons between OVERFLOW calculations and tests of two vehicles in the Langley Unitary Plan Wind Tunnel. Gridding and discretization, various measures of solution convergence, and turbulence modelling options are dicussed. While the presented guidelines were developed from OVERFLOW simulations in support of a particular test, the recommendations on numerical methods and gridding practices hold merit for simulations conducted CFD solvers in general.

CFD↗
NASA NTRSDate not supplied

Aeroacoustic Analysis using Dynamic Mode Decomposition of Unsteady Pressure-Sensitive Paint Measurements

In this paper, we present a new method for the diagnosis and analysis of aeroacoustic phenomena. Taking advantage of the well-known property that the measurements of the Unsteady Pressure-Sensitive Paint (uPSP) have much higher spatial resolution compared to those of the conventional pressure transducers, the method is based on the visualization and analysis of the outputs of the Dynamic Mode Decomposition (DMD) of the uPSP measurements. The uPSP measurements were collected with four Phantom high-speed cameras in the Ascent Transient Aerodynamics Test (ATAT) of the Space Launch System (SLS) Block 1 cargo vehicle in the 11-by-11-foot transonic test section of the Unitary Plan Wind Tunnel (UPWT) at NASA Ames Research Center in September 2019. The method presented in this paper is demonstrated by investigating an interesting phenomenon observed in the SLS ATAT – the vortex shedding tone generated by the forward attachment of the Solid Rocket Booster (SRB). As examples, the DMD outputs of the uPSP measurements in a subsonic test in the SLS ATAT are presented. It is shown the information retrieved from the DMD outputs of the uPSP measurements can be effectively used in the identification and diagnosis of the aeroacoustic phenomena. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding was provided by the NASA Aerosciences Evaluation and Test Capabilities Portfolio Office.

acoustics↗
NASA NTRSDate not supplied

Aeroacoustic Analysis using Dynamic Mode Decomposition of Unsteady Pressure-Sensitive Paint Measurements

In this paper, we present a new method for the diagnosis and analysis of aeroacoustic phenomena. Taking advantage of the well-known property that the measurements of the Unsteady Pressure-Sensitive Paint (uPSP) have much higher spatial resolution compared to those of the conventional pressure transducers, the method is based on the visualization and analysis of the outputs of the Dynamic Mode Decomposition (DMD) of the uPSP measurements. The uPSP measurements were collected with four Phantom high-speed cameras in the Ascent Transient Aerodynamics Test (ATAT) of the Space Launch System (SLS) Block 1 cargo vehicle in the 11-by-11-foot transonic test section of the Unitary Plan Wind Tunnel (UPWT) at NASA Ames Research Center in September 2019. The method presented in this paper is demonstrated by investigating an interesting phenomenon observed in the SLS ATAT – the vortex shedding tone generated by the forward attachment of the Solid Rocket Booster (SRB). As examples, the DMD outputs of the uPSP measurements in a subsonic test in the SLS ATAT are presented. It is shown the information retrieved from the DMD outputs of the uPSP measurements can be effectively used in the identification and diagnosis of the aeroacoustic phenomena. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding was provided by the NASA Aerosciences Evaluation and Test Capabilities Portfolio Office.

acoustics↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results of the CobraMRV Mars Entry Vehicle Concept during Supersonic Retropropulsion

The CFD solver Overflow was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a CFD and wind tunnel integration effort. Selection of numerical methods and resulting comparisons to selected tunnel conditions of interest are detailed. Comparisons of mean pressures with static taps and pressure-sensitive paint are provided for multiple SST-based turbulence models. URANS predicitons of heat shield pressures at high thrust conditions are poor compared to TMRC and DES predictions. Comparisons to experimental measurements of pressure fluctuations indicate complex relationships between heat shield unsteadiness, flight Mach number, and engine thrust coefficient. Bow shock topology and shock cell structure between experimental and OVERFLOW data show good agreement in Schlieren imaging. Recommendations for future investigation into flow field unsteadiness and turbulence models are suggested.

CFD↗
NASA NTRS1962

Unitary Wind Tunnel Tests of 30-Degree Conical Ribbon Parachute and a Rotofoil Parachute Towed in the Wake of a Conical Nosed Cylindrical Body

Multiple wind tunnel test trials were conducted on a 30 degree conical ribbon parachute with porosities of 30, 27, and 24 percent. Variables were Mach number, dynamic pressure, towline length, and coefficient of drag. A Rotofoil parachute having a porosity of approximately 24 percent was tested, but failed after about 30 seconds of operation at a Mach number of 1.8 All of the parachutes had a nominal diameter and shroud line length of 10 inches. Drag coefficients were based on the area of a circle having a diameter two-thirds of the nominal parachute diameter.

WIND TUNNEL TESTS↗
NASA NTRS2015

11 Foot Unitary Plan Tunnel Facility Optical Improvement Large Window Analysis

The test section of the 11 by 11-foot Unitary Plan Transonic Wind Tunnel (11-foot UPWT) may receive an upgrade of larger optical windows on both the North and South sides. These new larger windows will provide better access for optical imaging of test article flow phenomena including surface and off body flow characteristics. The installation of these new larger windows will likely produce a change to the aerodynamic characteristics of the flow in the Test Section. In an effort understand the effect of this change, a computational model was employed to predict the flows through the slotted walls, in the test section and around the model before and after the tunnel modification. This report documents the solid CAD model that was created and the inviscid computational analysis that was completed as a preliminary estimate of the effect of the changes.

computer aided design↗
NASA NTRS1999

A Computational/Experimental Study of Two Optimized Supersonic Transport Designs and the Reference H Baseline

Two supersonic transport configurations designed by use of non-linear aerodynamic optimization methods are compared with a linearly designed baseline configuration. One optimized configuration, designated Ames 7-04, was designed at NASA Ames Research Center using an Euler flow solver, and the other, designated Boeing W27, was designed at Boeing using a full-potential method. The two optimized configurations and the baseline were tested in the NASA Langley Unitary Plan Supersonic Wind Tunnel to evaluate the non-linear design optimization methodologies. In addition, the experimental results are compared with computational predictions for each of the three configurations from the Enter flow solver, AIRPLANE. The computational and experimental results both indicate moderate to substantial performance gains for the optimized configurations over the baseline configuration. The computed performance changes with and without diverters and nacelles were in excellent agreement with experiment for all three models. Comparisons of the computational and experimental cruise drag increments for the optimized configurations relative to the baseline show excellent agreement for the model designed by the Euler method, but poorer comparisons were found for the configuration designed by the full-potential code.

Cliff, Susan E.↗
NASA NTRS1974

Investigations on an 0.030-scale space shuttle vehicle configuration 140A/B orbiter model in the Ames Research Center unitary plan 8 by 7-foot supersonic wind tunnel (0A53C)

A wind tunnel test was conducted of an 0.030 scale model of the space shuttle orbiter in a supersonic wind tunnel. Tests were conducted at Mach numbers of 2.5, 3.0, and 3.5. Reynolds numbers ranged from 0.75 million per foot to 4.00 million per foot. The objective of the test was to establish and verify longitudinal and lateral-directional aerodynamic performance, stability, and control characteristics for the configuration 140 A/B SSV Orbiter. Six-component force and moment data, base and cavity pressures, body-flap, elevon, speedbrake, and rudder hinge moments, and vertical tail forces and moments were measured.

Nichols, M. E.↗
NASA NTRS2011

Use of Heated Helium to Simulate Surface Pressure Fluctuations on the Launch Abort Vehicle During Abort Motor Firing

The solid-rocket plumes from the Abort motor of the Multi-Purpose Crew Vehicle (MPCV, also know as Orion) were simulated using hot, high pressure, Helium gas to determine the surface pressure fluctuations on the vehicle in the event of an abort. About 80 different abort situations over a wide Mach number range, (0.3< or =M< or =1.2) and vehicle attitudes (+/-15deg) were simulated inside the NASA Ames Unitary Plan, 11-Foot Transonic Wind Tunnel. For each abort case, typically two different Helium plume and wind tunnel conditions were used to bracket different flow matching critera. This unique, yet cost-effective test used a custom-built hot Helium delivery system, and a 6% scale model of a part of the MPCV, known as the Launch Abort Vehicle. The test confirmed the very high level of pressure fluctuations on the surface of the vehicle expected during an abort. In general, the fluctuations were found to be dominated by the very near-field hydrodynamic fluctuations present in the plume shear-layer. The plumes were found to grow in size for aborts occurring at higher flight Mach number and altitude conditions. This led to an increase in the extent of impingement on the vehicle surfaces; however, unlike some initial expectations, the general trend was a decrease in the level of pressure fluctuations with increasing impingement. In general, the highest levels of fluctuations were found when the outer edges of the plume shear layers grazed the vehicle surface. At non-zero vehicle attitudes the surface pressure distributions were found to become very asymmetric. The data from these wind-tunnel simulations were compared against data collected from the recent Pad Abort 1 flight test. In spite of various differences between the transient flight situation and the steady-state wind tunnel simulations, the hot-Helium data were found to replicate the PA1 data fairly reasonably. The data gathered from this one-of-a-kind wind-tunnel test fills a gap in the manned-space programs, and will be used to establish the acoustic environment for vibro-acoustic qualification testing of the MPCV.

Panda, Jayanta↗
NASA NTRSDate not supplied

Flow Characterization of the NASA Langley Unitary Plan WindTunnel, Test Section 2: Computational Results

Flow in the empty Unitary Planform Wind Tunnel at the NASA Langley Research Center issimulated with computational fluid dynamics methods. The objectives are to assess CFD’s truepredictive capability and to generate flow-field maps upstream of the tunnel’s test section foruse in associated vehicle tests. Multiple CFD solvers, grid adaption methods, and turbulencemodels are used by five teams doing the simulations. The simulation domain is as large aspractical: from the start of the settling chamber, just downstream of the last set of turningvanes, to well downstream of the test section. Simulations are done at three Mach numbers andseveral Reynolds numbers that cover most of the tunnel’s operating range. The sensitivity ofthe CFD solutions to simulation process details, including turbulence model, spatial resolutionand settling-chamber inflow velocity profile, is characterized. Significant sensitivities in somedetails of the predicted test section flow are described. CFD predicts that details of flow inthe settling chamber can affect the flow in the test section. Turning of the flow through thetunnel contraction and past nozzle-block hardware generates streamwise vortices, some ofwhich survive into the core of the test section, principally at low Mach number. Preliminarycomparisons to experimental measurements are given. The major flow characteristics andthe dynamical vortices predicted by CFD exist in the experiment. Several details in CFD andexperiment are observed to differ, including the effects of the vortices and features that appearto be Mach waves in the test section. Sensitivity to grid resolution and turbulence modelingis noted in the vortices; no potential cause for the differences in Mach waves has yet beenidentified in the CFD sensitivity studies.

wind tunnel↗
NASA NTRS2015

Computational and Experimental Unsteady Pressures for Alternate SLS Booster Nose Shapes

Delayed Detached Eddy Simulation (DDES) predictions of the unsteady transonic flow about a Space Launch System (SLS) configuration were made with the Fully UNstructured Three-Dimensional (FUN3D) flow solver. The computational predictions were validated against results from a 2.5% model tested in the NASA Ames 11-Foot Transonic Unitary Plan Facility. The peak C(sub p,rms) value was under-predicted for the baseline, Mach 0.9 case, but the general trends of high C(sub p,rms) levels behind the forward attach hardware, reducing as one moves away both streamwise and circumferentially, were captured. Frequency of the peak power in power spectral density estimates was consistently under-predicted. Five alternate booster nose shapes were assessed, and several were shown to reduce the surface pressure fluctuations, both as predicted by the computations and verified by the wind tunnel results.

Braukmann, Gregory J.↗