Wind-Tunnel Sonic-Boom Testing Techniques
Wind tunnel tests of sonic boom phenomena, noting weak pressure field measurement, construction of extremely small models, boundary layer effects, etc
WIND TUNNEL↗
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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.
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Experimental and Computational Study of Convective Heating on the Aeroshell Back Side and Payload of the LOFTID Vehicle
A wind tunnel test program was conducted to obtain experimental convective heating data on models of the Low Earth Orbital Flight Test of an Inflatable Decelerator (LOFTID) vehicle in support of the flight test program. Wind tunnel testing was conducted at Mach 6 and Mach 10 at unit Reynolds numbers from 0.5×106/ft to 4.0×106/ft on 6.0inch diameter aeroshell models. Global heating data were obtained through phosphor thermography measurements on the front and back of the aeroshell and on the payload. For all test conditions, attached, laminar flow was definitively produced on the front of the aeroshell, while in the wake the flow was separated, unsteady and possibly transitional or turbulent. A complementary computational study was performed to assess the accuracy of heating predictions for LOFTID wake flow environments. A new grid topology was developed for the study to allow the use of structured grid solvers for complex, open-backed geometries such as that of LOFTID. Close agreement between measurements and predictions was obtained on the front of the aeroshell. Comparisons were more challenging in the wake because of both the extremely low heating rates and the accompanying high uncertainties and the unsteady nature of the separated wake flow simulations. While the differences between heating predictions and measurements for the wake flow were higher than for the forebody flow, these differences were used to assist in the definition of conservative uncertainty margins for predictions of the flight environments for the successful flight test of the LOFTID vehicle in November 2022.
LOFTID↗
Effects of forward velocity on noise for a J85 turbojet engine with multitube suppressor from wind tunnel and flight tests
Flight and wind tunnel noise tests were conducted using a J85 turbojet engine as a part of comprehensive programs to obtain an understanding of forward velocity effects on jet exhaust noise. Nozzle configurations of primary interest were a 104-tube suppressor with and without an acoustically-treated shroud. The installed configuration of the engine was as similar as possible in the flight and wind tunnel tests. Exact simultaneous matching of engine speed, exhaust velocity, and exhaust temperature was not possible, and the wind tunnel maximum Mach number was approximately 0.27, while the flight Mach number was approximately 0.37. The nominal jet velocity range was 450 to 640 m/sec. For both experiments, background noise limited the jet velocity range for which significant data could be obtained. In the present tests the observed directivity and forward velocity effects for the suppressor are more similar to predicted trends for internally-generated noise than unsuppressed jet noise.
Stone, J. R.↗
Flow Visualization to Complement Modern Wing Tunnel Testing
In recent times, wind tunnel testing at the NASA Ames Unitary Plan Wind Tunnel (UPWT) seeks to validate computational fluid dynamics (CFD) tools. To that end, we find that now more than ever, flow visualization is a powerful capability that provides context to the experimental results bringing real-time insight to the researcher. It enables us to answer questions like: “Why does this drag polar look different?” or “What is the source of this tone in the microphone data?”. Flow visualization can verify that assumptions made in the CFD model are correct, addressing questions like: “Are our trip dots the correct size and location?” or “Does the position of this shock match our prediction?”. Techniques such as infrared flow visualization and schlieren / shadowgraph are powerful capabilities that when taken together can provide results as to the nature of the fluid structure interaction at a test condition in a format that pairs well with computational results. When properly integrated these measurements can be taken with no impact on the productivity of the test program, supplementing traditional measurements and often discovering something unforeseen, helping the researcher to answer those slippery aerodynamic questions and ultimately build confidence in the accuracy of a model.
wind tunnel↗
Helicopter testing in a wind tunnel
Wind tunnel testing of helicopters, and test program for Polish helicopter industry
Brodzki, Z.↗
Correlation of flight-test loads with wind- tunnel predicted loads on three lifting body vehicles
An essential area of research with the unique M2-F2, HL-10, and X-24A lifting body configurations is the assessment of the ability to predict flight loads from wind- tunnel tests. Flight measurements and correlation with predictions are necessary in verifying the structural integrity of existing vehicles and establishing the groundwork for weight savings on future vehicles of similar shapes. As part of the overall lifting body flight investigation at the Flight Research Center, detailed aerodynamic-load studies are being made on each of the three vehicles. This paper presents the preliminary results from these studies
Ming H Tang↗
An Investigation of the McDonnell XP-85 Airplane in the Ames 40- by 80-foot Wind Tunnel. Force and Moment Tests
Wind-tunnel tests of the McDonnell XP-85 airplane were conducted to determine its longitudinal, lateral, and directional stability and the characteristics of the aileron, the ruddervator, the leading-edge droop nose flap, and the stall control vanes. The directional stability of the airplane with numerous skyhook modifications and with a ventral fin was also investigated. The results of the tests showed that the effectiveness of the droop nose flaps and the stall control vanes was negligible with regard to either the maximum lift or longitudinal stability of the airplane. Contrary to any previous small-scale results, extension of the skyhook caused a 75-percent reduction in the directional stability of the airplane for both low and high values of lift coefficient. The simplest solution to the problem short of a major redesign of the skyhook appears to be the adoption of a ventral fin.
Hunton, Lynn W.↗
Integrated Aeroservoelastic Optimization: Status and Direction
The interactions of lightweight flexible airframe structures, steady and unsteady aerodynamics, and wide-bandwidth active controls on modern airplanes lead to considerable multidisciplinary design challenges. More than 25 years of mathematical and numerical methods' development, numerous basic research studies, simulations and wind-tunnel tests of simple models, wind-tunnel tests of complex models of real airplanes, as well as flight tests of actively controlled airplanes, have all contributed to the accumulation of a substantial body of knowledge in the area of aeroservoelasticity. A number of analysis codes, with the capabilities to model real airplane systems under the assumptions of linearity, have been developed. Many tests have been conducted, and results were correlated with analytical predictions. A selective sample of references covering aeroservoelastic testing programs from the 1960s to the early 1980s, as well as more recent wind-tunnel test programs of real or realistic configurations, are included in the References section of this paper. An examination of references 20-29 will reveal that in the course of development (or later modification), of almost every modern airplane with a high authority active control system, there arose a need to face aeroservoelastic problems and aeroservoelastic design challenges.
Livne, Eli↗
Design integration and noise study for a large STOL augmentor wing transport
Analysis, design, experimental static testing, wind-tunnel testing, and design integration studies are used to develop an augmentor wing jet flap configuration for a jet STOL transport aircraft having maximum propulsion and aerodynamic performance with minimum noise generation. The program has three basic elements: (1) static testing of a scale wing section to demonstrate augmentor performance and noise characteristics; (2) two-dimensional wind-tunnel testing to determine flight-speed effects on performance, and (3) system design and evaluation that optimizes the complete system and ensures that the design is compatible with the requirements for a large STOL transport having a 500-foot sideline noise of 95 perceived noise decibels (PNdb) or less.
Okeeke, J. V.↗
Theoretical study of conditions limiting V/STOL testing in wind tunnels with solid floor
Conditions limiting V/STOL testing in wind tunnels with solid floors
Heyson, H. H.↗
Effects of forward velocity on noise for a J85 turbojet engine with multitube suppressor from wind tunnel and flight tests
Flight tests and wind tunnel noise tests using a J85 engine with some representative jet exhaust noise suppressors are reported. Capabilities and limitations of the two types of experimental methods are evaluated with emphasis on investigation of forward velocity effects. The suppressor arrangement was a 104-elliptical-tube nozzle configuration, with or without an acoustically lined shroud. The suppressor noise levels are found not reduced as much by forward velocity as expected for unsuppressed jets. The directivity and forward velocity effects appear more similar to predicted trends for internally generated noise than for unsuppressed jet noise.
Stone, J. R.↗
Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel
Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.
Mars↗
Flow Quality Survey of the 8- by 6-Foot Supersonic Wind Tunnel (2015 Test) Prior to the 9- by 15-Foot Acoustic Low-Speed Wind Tunnel Acoustic Improvement Modifications
Prior to the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel (9x15 LSWT) undergoing significant structural and acoustic modifications beginning in 2017, a flow quality survey (FQS) was conducted to characterize the flow quality of the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT) test section and facility air dryer. The data from this FQS will be compared to data collected during a FQS following the completion of the 9x15 LSWT Acoustic Improvement Modifications. This document contains only the FQS data collected in November 2015 using the 16-inch-diameter cone cylinder and a set of twenty air dryer bed wind anemometers. The ability to reach the extents of the 8x6 SWT operating envelope with a large blockage model was verified through the 16-inch-diameter cone cylinder tests. The flow uniformity and angularity at the entrance of the facility air dryer was assessed and quantified using the air dryer bed wind anemometers.
Flow Quality Survey↗
Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel
Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.
Supersonic Retropropulsion↗
Techniques in Assessing Random Uncertainty of Wind Tunnel Replicate Test Data
The preferred method to determine the random uncertainty of variables of interest during wind tunnel tests is by direct analysis of replicate data. However, current information is limited to analysis of a single replicate set condition that does not consider data covariance. This paper provides additional information on considerations for quantifying random uncertainty including the evaluation and correction of replicate data covariance, the selection of replicate set conditions for the test, and the interpolation of random uncertainty at other than replicate set conditions. The analysis and techniques are supported by data from a conventional aircraft configuration test at the Oran Nicks Low-Speed Wind Tunnel and an ASME nozzle flow characterization at the Propulsion Systems Laboratory.
Uncertainty↗
Evaluation of flow quality in two NASA transonic wind tunnels
Tests have been conducted in the Langley Research Center 8-foot Transonic Pressure Tunnel and the Ames Research Center 12-foot pressure wind tunnel in order to measure characteristic disturbance levels and energy spectra in their respective settling chambers, test sections, and diffusers and to determine the sources of these disturbances. Results are presented and discussed along with some specific recommendations.
Owen, F. K.↗
F/A-18 forebody vortex control. Volume 1: Static tests
A wind tunnel test was conducted on a six percent model of the F/A-18 at the NASA Ames 7 X 10-Foot Low Speed Wind Tunnel. The primary objective of the test was to evaluate several forebody vortex control configurations at high angles of attack in order to determine the most effective method of obtaining well behaved yawing moments, in preparation for the rotary balance test. Both mechanical and pneumatic systems were tested. Single and dual rotating nose tip strakes and a vertical nose strake were tested at different sizes and deflections. A series of jet blowing configurations were located at various fuselage stations, azimuth angles, and pointing angles ranging from straight aft to 60 deg canted inboard. Slot blowing was investigated for several slot lengths and fuselage stations. The effect of blowing rate was tested for both of these pneumatic systems. The most effective configurations were then further tested with a variation of both sideslip angle and Reynolds number over a range of angles of attack from 0 to 60 deg. It was found that a very robust system can be developed that provides yawing moments at angles of attack up to 60 deg that significantly exceeds that available from 30 deg of rudder deflection (F/A-18 maximum) at 0 deg angle of attack.
Kramer, Brian R.↗
Comparison of Wind-Tunnel and Flight Unsteady Pressure Stochastic Characteristics for the Space Launch System Artemis I Flight
Over the course of more than ten years, numerous wind-tunnel tests have been conducted to acquire data for characterizing the unsteady pressure environments expected to act on the Space Launch System Block 1 crew launch vehicle during ascent. These wind-tunnel tests of highly-instrumented rigid models are the current standard for the estimation of unsteady environments. Following the successful launch of the Artemis I mission, the extensive flight data acquired can be analyzed to evaluate the accuracy of unsteady pressure environments predicted in subscale wind-tunnel testing in comparison to the flight test data. In this paper, analyses focusing on data from several Space Launch System wind-tunnel tests and the Artemis I flight test are presented, including assessments of test-to-test, tunnel-to-tunnel, and tunnel-to-flight stochastic characteristics and preflight modeling validity based on wind-tunnel testing. In general, the fluctuating pressure environments measured during the several preflight subscale wind-tunnel tests compare favorably and provide relatively accurate estimates of the environments measured during flight. Discrepancies in fluctuating magnitudes, spatial extent of regions of unsteadiness, and narrowband frequency peaks are noted in the multibody region aft of the solid rocket booster forward attachment to the core stage.
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