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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Wind Tunnel Testing of Next-Generation High-Speed Engines

This talk will introduce students to supersonic and hypersonic flows as well as cover why these high-speed flow regimes are important to us today. The main focus is on the different propulsion systems used in Mach 3+ flight and how they are tested in wind tunnel facilities. Different wind tunnels at NASA Langley will used to demonstrate how ground tests facilitate the development of new engines from a proof-of-concept to a mature flight qualified design. There are many opportunities in field of high-speed aeronautics. I'll discuss my path to NASA Langley then give advice for internship and job opportunities at NASA.

Robin Hunt↗
NASA NTRSDate not supplied

Mars Sample Return, Sample Retrieval Lander, Reaction Control System Jet Interaction Supersonic Wind Tunnel Test Overview with CFD Predictions

NASA's Mars Sample Return campaign will be launching several missions over the next decade that will work together to return rock samples from Mars back to Earth. The Sample Retrieval Lander (SRL) will deliver the Mars Ascent Vehicle and fetch rover to the surface of Mars in 2006. Rock samples collected by the Mars 2020 Perseverance rover, landing in early 2021, will be loaded on the the ascent vehicle to be launched into Mars orbit for retrieval by yet another spacecraft. The Sample Retrieval Lander will be a blunt entry capsule similar to past Mars entry vehicles like Mars Science Laboratory and Viking. The vehicle will fly a guided entry, using a small lift vector produced by a non-zero trim angle of attack to eliminate downrange and crossrange position errors at the point of parachute deploy. This energy and heading management is achieved with a reaction control system (RCS) that directs the bank angle of the vehicle and also minimizes unwanted capsule dynamics. The reaction control system and control design is based on the Mars Science Laboratory and Mars 2020 RCS systems. However, due to packaging constraints, the backshell of this new entry vehicle has a different geometry than those earlier designs. To certify the RCS system for flight the project must characterize the jet plume interactions with the capsule backshell that could impair or significantly augment the RCS control authority. This characterization will be done through a combination of computational fluid dynamics (CFD) analysis and wind tunnel test. Two candidate arrangements of the RCS jets have been identified for the SRL vehicle and are currently under evaluation before final selection. The aero/RCS plume interactions of these candidate configurations have been measured in a supersonic wind tunnel test in NASA Langley's Unitary Plan Wind Tunnel. The test was conducted in the fall of 2020 and data is currently being reduced. An overview of the candidate RCS configurations are presented here with an overview of the wind tunnel model design, jet scaling and scaled nozzle design, and the test matrix. Preliminary CFD runs are presented with an assessment of the predicted plumes and their interaction with the wake flow of the vehicle. The predicted effects of the model sting is provided as well. This high fidelity wind tunnel test is being conducted much earlier in the SRL project than would normally be done. The test was funded as part of a CFD evaluation task funded by NASA's Aerosciences Evaluation and Test Capabilities Project. The objective of the evaluation task was to compare the ability of CFD to predict complex flows with data that can be measured in the Langley Unitary Plan Wind Tunnel. RCS Jet interactions were selected as a type of complex flow that is important to NASA missions. In addition to providing useful data to the SRL project, there was added emphasis on quantifying the accuracy of the CFD predictions and wind tunnel test data. An overview of the uncertainty quantification methodologies for computational and experimental portions of this test is presented.

blunt body↗
NASA NTRSDate not supplied

Aero-Propulsive Damping Characterization for eVTOL Aircraft Using Free Motion Wind-Tunnel Testing

This paper describes an electric vertical takeoff and landing (eVTOL) aircraft system identification method applied using three degree-of-freedom (3DOF) free motion wind-tunnel testing. The approach, similar to flight-test system identification, allows for efficient mathematical model development of the aero-propulsive moments applied on an eVTOL vehicle, including aerodynamic damping effects. The approach is demonstrated using a subscale tiltrotor eVTOL aircraft mounted on a new 3DOF wind-tunnel apparatus. To execute the test, a model-based 3DOF control system is designed to track attitude commands and transition the aircraft based on the freestream dynamic pressure. While the flight controller is active, orthogonal phase-optimized multisine inputs are injected into the attitude command and control effector command signals to enable collection of informative data for model identification. Aero-propulsive models are then identified at several reference conditions in the transition flight envelope using the equation-error method in the frequency domain. The identified models are shown to have a good fit to the modeling data and good prediction capability of data not used for model identification. The method yields aerodynamic damping estimates using less wind-tunnel test time compared to traditional forced oscillation experiments and supplements static wind-tunnel testing to produce a comprehensive transition aero-propulsive model suitable for use in flight dynamics simulations.

RAVEN↗
NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗
NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗
NASA NTRS1982

Investigation of correlation between full-scale and fifth-scale wind tunnel tests of a Bell helicopter Textron Model 222

Reasons for lack of correlation between data from a fifth-scale wind tunnel test of the Bell Helicopter Textron Model 222 and a full-scale test of the model 222 prototype in the NASA Ames 40-by 80-foot tunnel were investigated. This investigation centered around a carefully designed fifth-scale wind tunnel test of an accurately contoured model of the Model 222 prototype mounted on a replica of the full-scale mounting system. The improvement in correlation for drag characteristics in pitch and yaw with the fifth-scale model mounted on the replica system is shown. Interference between the model and mounting system was identified as a significant effect and was concluded to be a primary cause of the lack of correlation in the earlier tests.

Squires, P. K.↗
NASA NTRS1979

Design Guidelines for the Application of Forebody and Nose Strakes to a Fighter Aircraft Based on F-16 Wind Tunnel Testing Experience

During the YF-16 and F-16 developmental wind tunnel test program, numerous variations in nose and forebody strakes were investigated. These data were reviewed, and the strake aerodynamic characteristics coalesced into design guidelines for the application of strakes to fighter aircraft. The design guides take the form of general equations governing the modification of forebody strakes to obtain a linear pitching moment curve and the calculation of the resulting lift and drag increments. Additionally, qualitative comments are made concerning the effects of strake geometry on lateral/directional stability. It is concluded that the generation of incremental strake lift is primarily dependent upon the area affected by the strake-induced vortex and that strake planform is of secondary importance. Forebody strakes have small beneficial effects on lateral/directional stability if properly designed; however, significant gains are easily attained with nose strakes.

Smith, C. W.↗
NASA NTRSDate not supplied

Aero-Propulsive Modeling for eVTOL Aircraft Using Wind Tunnel Testing with Multisine Inputs

A novel approach for modeling the aero-propulsive characteristics of an electric vertical takeoff and landing (eVTOL) aircraft was developed and demonstrated in wind tunnel testing. The approach was applied to the NASA LA-8 tandem tilt-wing eVTOL aircraft, using an efficient, hybrid experiment design composed of a static I-optimal response surface design for slowly-varying test variables, and dynamic orthogonal phase-optimized multisine excitations for the control surfaces and electric propulsors. Both the static and dynamic experiment designs were executed simultaneously to collect informative data for model identification. Statistically-weighted multivariate orthogonal function modeling was used to combine local modeling results computed in the frequency domain using data collected with dynamic excitations operating on the control effectors to form an aggregate aero-propulsive model. The final identified model exhibited good predictive capability when compared to validation data acquired separately from the data used to develop the model. The required test time using these new techniques was reduced by at least a factor of five compared to previous static wind tunnel testing for the LA-8 aircraft, while providing more informative data, greater parameterization flexibility, and high-quality models.

system identification↗
NASA NTRS1987

Wind tunnel tests on a one-foot diameter SR-7L propfan model

Wind tunnel tests have been conducted on a one-foot diameter model of the SR-7L propfan in the Langley 16-Foot and 4 x 7 Meter Wind Tunnels as part of the Propfan Test Assessment (PTA) Program. The model propfan was sized to be used on a 1/9-scale model of the PTA testbed aircraft. The model propeller was tested in isolation and wing-mounted on the aircraft configuration at various Mach numbers and blade pitch angles. Agreement between data obtained from these tests and data from Hamilton Standard validate that the 1/9-scale propeller accurately simulates the aerodynamics of the SR-7L propfan. Predictions from an analytical computer program are presented and show good agreement with the experimental data.

Aljabri, Abdullah S.↗
NASA NTRS1992

On-line analysis capabilities developed to support the AFW wind-tunnel tests

A variety of on-line analysis tools were developed to support two Active Flexible Wing wind-tunnel tests. These tools were developed to verify control law execution, to satisfy analysis requirements of the control law designers, to provide measures of system stability in a real-time environment, and to provide project managers with a quantitative measure of controller performance. Description and purposes of capabilities which were developed are presented in this paper along with examples. Procedures for saving and transferring data for near real-time analysis, and descriptions of the corresponding data interface programs are also presented. The on-line analysis tools worked well before, during, and after the wind-tunnel tests and proved to be a vital and important part of the entire test effort.

Wieseman, Carol D.↗
NASA NTRS1974

Wind tunnel test of the 0.019 (2A configuration) jet plume space shuttle integrated vehicle in the ARC 9- by 7-foot unitary wind tunnel (IA12B)

The wind tunnel test of the 0.019 jet plume space shuttle integrated vehicle in the Ames 9 ft by 7 ft unitary wind tunnel was conducted at Mach numbers of 1.55 and 2.0 over a Reynolds number range from 3.5 million to 4.1 million/ft. Data were obtained at angles of attack from minus 8 deg to plus 8 deg at 0 deg sideslip and at angles of sideslip from minus 9 deg to plus 8 deg at 0 deg angle of attack. The basic configuration tested was the 2A vehicle with the orbiter at 0 deg angle of incidence with respect to the external tank. The other deviations to the 2A configuration were the solid rocket motor shrouds, which were designed to vehicle '3' lines, and the tank nose, which consisted of the retro-package being removed and replaced by a 16.5 inch full scale radius nose.

Hardin, R. B.↗
NASA NTRS2019

Additional Findings from the Common Research Model Natural Laminar Flow Wind Tunnel Test

An experimental investigation of the Common Research Model with Natural Laminar Flow (CRM-NLF) took place in the National Transonic Facility (NTF) at the NASA Langley Research Center in 2018. The 5.2% scale semispan model was designed using a new natural laminar flow design method, Crossflow Attenuated NLF (CATNLF). CATNLF enables laminar flow on typical transport wings with high sweep and Reynolds number by reshaping the wing airfoils to obtain specific pressure distribution characteristics that control the crossflow growth near the leading edge. The CATNLF method also addresses Tollmien- Schlichting transition, attachment line transition, and Görtler vortices. During the wind tunnel test, data were acquired to address three primary test objectives: validate the CATNLF design method, characterize the NTF laminar flow testing capabilities, and establish best practices for laminar flow wind tunnel testing. The present paper provides both experimental and computational data to understand the CRM-NLF laminar flow characteristics, as well as address the three primary test objectives. The effects of angle of attack and Reynolds number on the CRM-NLF laminar flow extent are studied, and the dominant transition mechanism is evaluated at a variety of test conditions. Critical N-factors are calculated for the NTF environment, and a discussion on best practices for laminar flow wind tunnel testing is provided. The CRM-NLF in the NTF provided initial confirmation of the ability of the CATNLF method to suppress crossflow growth and enable significant extents of laminar flow on transport wings with high sweep and Reynolds numbers.

Lynde, Michelle N.↗
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary layer flow interaction with the microphone holder seems to have some effect on the cavity resonance response, the NASA/Marshall Space Flight Center (MSFC) aeroacoustic team postulated that changing the design of the holder might reduce the cavity resonance response. An experiment was performed at the MSFC Trisonic Wind Tunnel comparing a new holder design with the traditional MSFC holder. The new holder design is intended to mimic the Kulite Semiconductors, Inc. B-screen, where the single hole of the traditional MSFC holder is replaced by a series of ten 0.007 inch diameter holes. This particular design serves to protect the transducer diaphragm as in the manufacturer's B-screen, but also increases the frequency of the XMIN such that it is not apparent in the data collected from aeroacoustic wind tunnel tests. MSFC compared the two holder designs by installing two transducers in standard holders on one side of a ten inch long ten degree cone, and two transducers in B-screen holders on the opposite side for a series of runs at different conditions. Over 11-13 February, 2013, the model was run at Mach numbers between 0.80 and 1.96 at angles of attack at 0 deg, plus or minus 1 deg, plus or minus 2 deg, plus or minus 4 deg, and plus or minus 8 deg. Model roll attitudes at 0 deg, 90 deg, and 180 deg were tested to isolate potential tunnel bias. This paper presents results from the test showing that the proposed holder design significantly reduces the influence of XMIN on measured fluctuating pressure levels.

Herron, A. J.↗
NASA NTRS1992

On-line analysis capabilities developed to support the AFW wind-tunnel tests

A variety of on-line analysis tools were developed to support two active flexible wing (AFW) wind-tunnel tests. These tools were developed to verify control law execution, to satisfy analysis requirements of the control law designers, to provide measures of system stability in a real-time environment, and to provide project managers with a quantitative measure of controller performance. Descriptions and purposes of the developed capabilities are presented along with examples. Procedures for saving and transferring data for near real-time analysis, and descriptions of the corresponding data interface programs are also presented. The on-line analysis tools worked well before, during, and after the wind tunnel test and proved to be a vital and important part of the entire test effort.

Wieseman, Carol D.↗