NASA Sonic Boom Wind Tunnel Testing
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Low-Boom Technology↗
Engineering topics
Patrick R Shea
Publications and source records attributed to Patrick R Shea.
Uncertainty Quantification of Artemis I Space Launch System Integrated Aerodynamics Databases
Accurate prediction of integrated aerodynamic forces and moments is a necessary part of aerospace vehicle development. This accuracy can be quantified in the form of an uncertainty model, which makes the prediction more useful within an integrated vehicle design effort. Aerodynamic force and moment databases were constructed for the Artemis~I mission of the Space Launch System vehicle. These databases reconcile data from multiple sources to yield unified predictions of how NASA's most advanced launch vehicle interacts with Earth's atmosphere as it ascends into orbit. This paper outlines how the uncertainty quantification was performed for these databases to ensure comprehensive and tractable uncertainty source coverage.
Michael W Lee↗
X-59 Sonic Boom Test Results from the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel
A wind tunnel test was conducted to investigate near-field sonic boom pressure signatures of the X-59 Low-Boom Flight Demonstrator aircraft. A 1.62%-scale model of the aircraft was fabricated for the wind tunnel test, which took place in the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel in September and October 2021. The model had provisions for being mounted by a swept blade strut that attached at top of model ahead of the inlet, or by a rear-entry sting that held the model at the location of the nacelle. The model had alternate parts for ±0.5° deflections of the flaps, ailerons, and stabilator, and ±1° deflections of the T-tail. Off-body static pressure measurements of the flow field below the model were made on a pressure rail which had 420 orifices along its tip. The model was positioned at various heights from the rail by vertical movement of the tunnel strut, and at various longitudinal stations relative to the rail by means of a linear actuator mounted between the wind tunnel strut and the balance. Spatial averaging of model pressure signatures acquired over a range of longitudinal positions reduced the effects of tunnel flow distortions and the interference of the rail flow field and shocks on the model pressure signatures. The test was run at approximate Mach numbers of 1.36, 1.4, and 1.47, and the model was set at various angles of attack and roll relative to the rail. Plots of the model signatures for representative variations of Mach number, model angles, control deflections, and height relative to the rail are provided throughout the report. Repeatability was generally very good and gave confidence in the quality of the measurements. The signatures measured at various heights from the rail provided insight into the aging of the model shocks as they propagated from 1.2 to 3 body lengths from the model. Off-track signatures up to 45° from centerline obtained by rolling the model gave indications of the shock flow fields across the width of the sonic boom carpet. The deflections of the various control surfaces allowed assessment of the boom sensitivity to the control surface movements.
Sonic boom↗
Experimental Characterization of the Space Launch System Block 1B Liftoff and Transition Environment
A wind tunnel test was conducted at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel to evaluate the liftoff and transition flow environment of the Space Launch System Block 1B launch vehicles and tower interference effects with the newly designed Mobile Launcher 2. This test leveraged several unique diagnostic capabilities, including eight distributed force and moment measurements on the launch tower tiers, new umbilical and crew access arm configurations, and unsteady pressure acquisition on the rocket nose in the cargo configuration. The tower loading revealed a strong impact of the SLS vehicle height on force coefficients for individual tiers. Although the umbilicals produced weak influence on the loading on both vehicle and launch tower, unsteady pressure measurements revealed a slight increase in fundamental pressure oscillation frequency at specific wind azimuth directions. Multimodal flow states were observed in the gap flow between the vehicle and solid rocket boosters, producing different load profiles along the vehicle that are now incorporated in the vehicle database. This phenomenon was investigated using time-series measurements of forces as well as tufts and smoke flow visualization.
Lee J Mears↗
Experimental Characterization of the Space Launch System Block 1B Liftoff and Transition Environment
A wind tunnel test was conducted at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel to evaluate the liftoff and transition flow environment of the Space Launch System Block 1B launch vehicles and tower interference effects with the newly designed Mobile Launcher 2. This test leveraged several unique diagnostic capabilities, including eight distributed force and moment measurements on the launch tower tiers, new umbilical and crew access arm configurations, and unsteady pressure acquisition on the rocket nose in the cargo configuration. The tower loading revealed a strong impact of the SLS vehicle height on force coefficients for individual tiers. Although the umbilicals produced weak influence on the loading on both vehicle and launch tower, unsteady pressure measurements revealed a slight increase in fundamental pressure oscillation frequency at specific wind azimuth directions. Multimodal flow states were observed in the gap flow between the vehicle and solid rocket boosters, producing different load profiles along the vehicle that are now incorporated in the vehicle database. This phenomenon was investigated using time-series measurements of forces as well as tufts and smoke flow visualization.
SLS↗
Aeroacoustic Computations of a Generic Low Boom Concept in Landing Configuration: Part 3 - Aerodynamic Validation and Noise Source Identification
Wind tunnel test data were used to validate the predicted aerodynamic behavior of a 15%-scale version of a generic, low-boom aircraft. The test was conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of the model. Measured steady surface pressures and global forces were used to validate predicted aerodynamic results obtained from high-fidelity simulations of the model as installed in the tunnel. Very good agreement between predicted and measured aerodynamic trends was demonstrated, providing the impetus to proceed with companion airframe noise simulations that were conducted in a free-air setting. Computed near-field flow variables acquired on a permeable data surface were used to generate synthetic pressure records for two 800-element phased microphone arrays positioned overhead and to the side of the model. The array data were beamformed to generate noise source localization maps for the clean model and several landing configurations. Primary and secondary airframe sources were identified and their relative strengths were determined. Far-field integrated noise spectra for the full aircraft, as well as individual components, were obtained from the source maps via integration of tailored regions. The analysis showed that noise produced by the landing gear was the dominant contributor to the far-field acoustic signature of the model, followed by flap noise. The effects of permeable data surface end caps, spatial resolution, array orientation, angle of attack, component interaction, velocity scaling, and numerical precision on synthetic far-field spectra were also evaluated.
low boom↗
Aeroacoustic Computations of a Generic Low Boom Concept in Landing Configuration: Part 3 - Aerodynamic Validation and Noise Source Identification
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low boom↗
Experimental Identification of Bistable Flow States on the Space Launch System at Liftoff Conditions
Time-averaged global force and moment measurements in wind tunnel tests can obscure complex flow dynamics, including bistability phenomena. This paper presents a detailed analysis into the potential cause of previously unobserved discrepancies among repeat runs during the 202114- by 22-Foot Subsonic Tunnel test of the Space Launch System. During the test, it was observed that liftoff configuration repeatability was not as strong as expected when compared to previous tests. These discrepancies occurred between wind angle azimuths of approximately 150◦ and 210◦, as well as between 330◦ and 30◦. After ruling out common confounding factors such as instrumentation hysteresis and temperature, it was hypothesized that the lack of repeatability was caused by the Coandă effect producing a bistable flow state at different locations between the cylindrical rocket centerbody and solid rocket boosters during the acquisition of data. Evidence to support this hypothesis is provided in this paper in the form of time-averaged force and moment coefficient data, time-dependent force and moment coefficient data, and still images from smoke flow and tuft flow visualization runs.
SLS↗
Unsteady Surface Pressures on the Space Launch System Block 1B Cargo Liftoff Configuration
A 1.75%-scale model of the NASA Space Launch System Block 1B Cargo vehicle was tested at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel. These experiments provided a test bed for validating unsteady surface pressure measurements acquired via lengths of pressure tubing and an ESP pressure scanner. Pressure was measured using an azimuthal array of pressure taps located on the nose of the Block 1B Cargo vehicle. Unique flow physics were observed in the wind tunnel test, especially the interaction between the launch vehicle and the support tower, which produced strong oscillations that were absent in the vehicle-alone configuration. It is suspected that a wake interaction produced resonant flow response when the tower is immediately downstream of the vehicle, and coherent pressure fluctuations were observed at all pressure taps in the azimuthal array.
Lee J Mears↗
Evaluating the Utility of Pressure Scanners for Unsteady Pressure Measurements in Wind Tunnel Characterization of the Space Launch System
In many aerodynamic measurement applications, the ability to make time-accurate measurements with pressure scanners provides practical advantages over making the same measurements directly using flush-mounted transducers. The pneumatic distortion due to the tubing between the sensing module and surface port limits the usefulness of such measurements to averaged, steady data only. This issue has restricted the application of pressure scanners for unsteady pressure measurements in wind tunnel and flight-testing applications. The Wiener-filtered inverse system response model has been proven to be effective in reconstructing complex unsteady pressure signals acquired remotely using pressure scanners. In this work, the feasibility of using pressure scanning systems for time-resolved pressure measurements in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel was evaluated. Results demonstrated good agreement between the reference transducers and reconstructed pressure scanner measurements up to frequencies of approximately 500 Hz.
Unsteady pressure↗
Improved Techniques for Measuring Static Ground Wind Loads on the NASA Space Launch System Mobile Launcher 2
Vertical distribution of static ground wind loads was directly measured on Mobile Launcher ground support hardware to the NASA Space Launch System (SLS) at 1.75% scale in the Langley Research Center 14- by 22-Foot Subsonic Tunnel. Low-speed wind tunnel testing in the liftoff phase of flight was conducted for the SLS Block 1B family of rockets with their associated ground support systems. Improved methods for measuring and estimating static ground wind loads on the launch tower assembly are introduced in detail. A novel, distributed loads measurement technique is introduced and compared against lower-fidelity, previously utilized methods.
Ground Wind Loads↗
Overview of the High Reynolds Number Ascent Wind Tunnel Test of the Space Launch System at the National Transonic Facility
Transonic, high Reynolds number wind tunnel testing for the Space Launch System (SLS) ascent flight environment was conducted in the National Transonic Facility (NTF) at the NASA Langley Research Center from December 2019 to April 2021. The test was sponsored by the SLS Program and the NASA Engineering & Safety Center with the ob- jective of assessing and characterizing Reynolds number effects on the ascent aerodynamics of the SLS launch vehicle. The cryogenic test article was a 1.75%-scale representation of the SLS Block 1 Cargo configuration that was fabricated mostly out of stainless steel, but with several additive-manufactured components including the Solid Rocket Booster (SRB) attach brackets and the SRB nozzles. This was the first use of additive manufacturing on structural components of a cryogenic model in the NTF, therefore post-fabrication material testing and inspections were performed to satisfy safety requirements. Force & moment and surface pressure data were acquired for Mach numbers between 0.50 and 0.95 over a range of Reynolds numbers based on core stage diameter between 2x10^6 and a maximum of 40x10^6 corresponding to 45% of flight Reynolds number. Additionally, flow visualization data using Pressure Sensitive Paint were acquired with a focus on the SRB forward attach area, which is susceptible to Reynolds number sensitivity at transonic speeds. This pa- per provides an overview of the test campaign including details on the unique test article and the experimental setup and test execution. General test findings and observations are also presented, but the majority of the test results and data analyses are provided in a companion paper.
Space Launch System↗
Improved Techniques for Measuring Static Ground Wind Loads on the NASA Space Launch System Mobile Launcher 2
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SLS↗
Force and Moment Analysis for the High Reynolds Number Wind Tunnel Test of the Space Launch System at Ascent Conditions
A high Reynolds number test of the Space Launch System was performed at the NASA Langley National Transonic Facility (NTF). The objective of the test was to use the cryogenic testing capabilities of the NTF to acquire data over the largest range of Reynolds numbers possible with a specific focus on the Reynolds numbers closest to flight conditions. The test was performed at Mach numbers from 0.50 to 0.95 which corresponds to the ascent portion of flight for the SLS vehicle. Force and moment data showed that pitching and yawing moment were sensitive to Reynolds number effects over the full range of Mach and Reynolds numbers tested. Axial force also showed sensitivity to Reynolds number with the largest differences seen between Mach 0.50 to 0.90. Surface pressure data showed the highest sensitivity to Reynolds number in the vicinity of the solid rocket booster forward attach region.
Space Launch System↗
Evaluation of Artemis I Aerodynamic Force and Moment Reconstructions and Pre-Flight Predictions for Ascent Flight
The Artemis I launch of the Space Launch System provided the first flight data for the new launch vehicle. The current paper shows comparisons of the preflight force and moment coefficient databases developed using traditional ground test and computation fluid dynamics to post-flight data reconstructions of the ascent aerodynamic force and moment coefficients. The post-flight reconstructions were developed using instrumentation on the vehicle, meteorological data, and Newton’s laws. In general, the preflight databases and post-flight reconstructed data showed similar trends throughout ascent from Mach 0.20 to Mach 3.50. Axial force coefficient showed the largest discrepancies, which is likely due to the difficulty of computing axial force from the flight data.
Space Launch System↗