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All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Model Deformation Measurement Technique NASA Langley HSR Experiences

Model deformation measurement techniques have been investigated and developed at NASA's Langley Research Center. The current technique is based upon a single video camera photogrammetric determination of two dimensional coordinates of wing targets with a fixed (and known) third dimensional coordinate, namely the spanwise location. Variations of this technique have been used to measure wing twist and bending at a few selected spanwise locations near the wing tip on HSR models at the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel. Automated measurements have been made at both the Transonic Dynamics Tunnel and at Unitary Plan Wind Tunnel during the past year. Automated measurements were made for the first time at the NTF during the recently completed HSR Reference H Test 78 in early 1996. A major problem in automation for the NTF has been the need for high contrast targets which do not exceed the stringent surface finish requirements. The advantages and limitations (including targeting) of the technique as well as the rationale for selection of this particular technique are discussed. Wing twist examples from the HSR Reference H model are presented to illustrate the run-to-run and test-to-test repeatability of the technique in air mode at the NTF. Examples of wing twist in cryogenic nitrogen mode at the NTF are also presented.

Burner, A. W.↗
NASA NTRS2020

Investigation of Atmospheric Boundary-Layer Effects on Launch-Vehicle Ground Wind Loads

A launch vehicle ground-wind-loads program is underway at the NASA Langley Transonic Dynamics Tunnel. The objectives are to quantify key aerodynamic and structural characteristics that impact the occurrence of large wind-induced oscillations of a launch vehicle when exposed to ground winds prior to launch. Of particular interest is the dynamic response of a launch vehicle when a von Kármán vortex street forms in the wake of the vehicle resulting in quasiperiodic lift and drag forces. Vehicle response to these quasiperiodic forces can become quite large when the frequency of vortex shedding nears that of a lowly-damped structural mode thereby exciting a resonant response. Wind approaching the vehicle can be characterized by a varying speed with height and turbulence content. The combination of both the varying speed and turbulence content is referred to herein as the atmospheric boundary-layer. The importance of the atmospheric boundary-layer upon launch vehicle wind-induced oscillation response has long been questioned, and its effects are not well understood. Although there are several facilities around the world dedicated to replicating atmospheric boundary layers, the development of such a boundary layer in a wind tunnel capable of producing flight-representative Reynolds numbers for aeroelastically-scaled launch vehicle models has only recently been accomplished. The NASA Langley Transonic Dynamics Tunnel is capable of simulating flight-representative Reynolds numbers of launch vehicles on the pad and is uniquely capable of replicating many fluid-structure scaling parameters typical of aeroelastic tests. Recent test efforts successfully developed representative atmospheric boundary-layers for three launch sites in the Transonic Dynamics Tunnel, thereby allowing all known aerodynamic and fluid-structure coupling parameters to be simultaneously simulated for those sites. Dynamic aeroelastically-scaled models representative of typical large launch vehicles were constructed for testing. Aeroelastic scaling includes matching geometry, mode shapes, reduced frequencies, damping, running mass ratios, and running stiffness ratios. The models were tested in smooth uniform flow and then immersed in the atmospheric boundary-layer for comparison of these effects. Dynamic data were acquired measuring unsteady pressure, acceleration, and base bending moment. It was discovered that peak dynamic loads resulting from resonant wind-induced oscillation response are similar when acquired in either smooth uniform flow or an atmospheric boundary-layer. This indicates that resonant lock-in events are minimally impacted by representative turbulence and/or wind profile. Alternately, nonresonant wind-induced oscillation response events are stronger when acquired in an atmospheric boundary-layer. This indicates that a lowly-damped structural response will increase when exposed to an increased magnitude of random excitation, which is consistent with historical comparisons. Loads created by the resonant response events were substantially stronger than those from the nonresonant response events. Therefore, if testing is done to simply identify worst-case conditions and load magnitude, then smooth uniform flow is likely an adequate test technique. However, if nonresonant response loads are of primary interest, then atmospheric boundary-layer simulation is required.

Thomas G Ivanco↗
NASA NTRS2015

Initial Assesment of Space Launch System Transonic Unsteady Pressure Environment

A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.

Sekula, Martin K.↗
NASA NTRS1986

Current wind tunnel capability and planned improvements at the NASA Langley Research Center

NASA Langley's major wing wind tunnels and the projected facilities planned to be completed by 1988 are presented. Special capabilities, uses, and the improvements done during the Langley's tunnel revitalization program are described for the following facilities: (1) the 30 x 60-ft subsonic tunnel, (2) the 4 x 7-m low-speed testing tunnel, (3) the Low-Turbulence Tunnel, (4) the Spin Tunnel, (5) the National Transonic Pressure Tunnel, (6) Transonic Cryogenic Tunnel, (7) the National Transonic Facility, (8) the 16-ft Transonic Tunnel, (9) the Transonic Dynamic Tunnel, (10) the Unitary Plan Wind Tunnel, and (11) a new 20-inch supersonic wind tunnel which is currently undergoing final checkout. The design concept of an extremely-low disturbance level supersonic tunnel, the upgrading plans for the hypersonic aerothermal complex, and the present and planned capabilities for testing the hydrogen-fueled Scramjet engines are also presented. In addition, uses of inexpensive simple-to-operate research wind tunnels are discussed. Tunnel diagrams and graphs of upgrade results are included.

Bower, R. E.↗
NASA NTRS2015

Time Accurate Unsteady Pressure Loads Simulated for the Space Launch System at a Wind Tunnel Condition

Using the Fully Unstructured Three-Dimensional (FUN3D) computational fluid dynamics code, an unsteady, time-accurate flow field about a Space Launch System configuration was simulated at a transonic wind tunnel condition (Mach = 0.9). Delayed detached eddy simulation combined with Reynolds Averaged Naiver-Stokes and a Spallart-Almaras turbulence model were employed for the simulation. Second order accurate time evolution scheme was used to simulate the flow field, with a minimum of 0.2 seconds of simulated time to as much as 1.4 seconds. Data was collected at 480 pressure taps at locations, 139 of which matched a 3% wind tunnel model, tested in the Transonic Dynamic Tunnel (TDT) facility at NASA Langley Research Center. Comparisons between computation and experiment showed agreement within 5% in terms of location for peak RMS levels, and 20% for frequency and magnitude of power spectral densities. Grid resolution and time step sensitivity studies were performed to identify methods for improved accuracy comparisons to wind tunnel data. With limited computational resources, accurate trends for reduced vibratory loads on the vehicle were observed. Exploratory methods such as determining minimized computed errors based on CFL number and sub-iterations, as well as evaluating frequency content of the unsteady pressures and evaluation of oscillatory shock structures were used in this study to enhance computational efficiency and solution accuracy. These techniques enabled development of a set of best practices, for the evaluation of future flight vehicle designs in terms of vibratory loads.

Alter, Stephen J.↗
NASA NTRSDate not supplied

Design of a Collocation-Based Active Flutter Suppression Control Law for the IAWTM Wind Tunnel Model

The design of an active flutter suppression (AFS) control law for upcoming wind tunnel tests in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center (LaRC) with the Integrated Adaptive Wing Technology Maturation (IAWTM) project is presented. The test article is a highly flexible half-span model of a transport airplane and tests will focus on the transonic regime. The control law is based on the concept of collocation, sometimes called identically located accelerometer and force (ILAF), which uses local velocity feedback to increase damping for all aeroelastic modes. A multiple-input multiple-output (MIMO) extension of this architecture is used for performance and robustness improvements. Results showed that the proposed control law with three fixed gains stabilized the design models over most of the test envelope and successfully extended the open-loop flutter boundary to higher Mach numbers and dynamic pressures.

Active flutter suppression↗
NASA NTRS1997

The Benchmark Active Controls Technology Model Aerodynamic Data

The Benchmark Active Controls Technology (BACT) model is a part of the Benchmark Models Program (BMP). The BMP is a NASA Langley Research Center program that includes a series of models which were used to study different aeroelastic phenomena and to validate computational fluid dynamics codes. The primary objective of BACT testing was to obtain steady and unsteady loads, accelerations, and aerodynamic pressures due to control surface activity in order to calibrate unsteady CFD codes and active control design tools. Three wind-tunnel tests in the Transonic Dynamics Tunnel (TDT) have been completed. The first and parts of the second and third tests focused on collecting open-loop data to define the model's aeroservoelastic characteristics, including the flutter boundary across the Mach range. It is this data that is being presented in this paper. An extensive database of over 3000 data sets was obtained. This database includes steady and unsteady control surface effectiveness data, including pressure distributions, control surface hinge moments, and overall model loads due to deflections of a trailing edge control surface and upper and lower surface

Scott, Robert C.↗
NASA NTRS2008

Aeroservoelastic Testing of a Sidewall Mounted Free Flying Wind-Tunnel Model

A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three j wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the rst of these three tests, a semispan, aeroelastically scaled, wind-tunnel model of a ying wing SensorCraft vehi- cle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid-body modes. Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Scott, Robert C.↗
NASA NTRS2011

Aeroservoelastic Wind-Tunnel Tests of a Free-Flying, Joined-Wing SensorCraft Model for Gust Load Alleviation

A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind-tunnel model of a joined-wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind-tunnel model was mated to a new, two-degree-of-freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at -10% static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free ying wind-tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

Scott, Robert C.↗
NASA NTRS2013

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression

of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a semispan, aeroelastically scaled, wind tunnel model of a flying wing SensorCraft vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Scott, Robert C.↗
NASA NTRS2014

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation

This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

Scott, Robert C.↗
NASA NTRSDate not supplied

Gust Load Alleviation Control and Gust Estimation for a High Aspect Ratio Wing Wind Tunnel Model

This paper presents the gust load alleviation (GLA) study of the aspect ratio 13.5 Common Research Model (CRM) wind-tunnel model. This study details the design of the GLA controller in preparation for wind-tunnel testing in the Transonic Dynamics Tunnel at NASA Langley Research Center. An aeroservolastic (ASE) model was first reduced using a model reduction method that takes advantage of the sinusoidal steady-state response. Then, the reduced model was used to design an extended-state Kalman filter which estimates the states and the sinusoidal gust input. The GLA control was then derived using the optimal control solution to a multi-objective cost function. The results of the GLA controller indicate a 75.6% reduction in wing-root strain while maintaining robust stability margins. The final paper will include performance under sub-optimal conditions, e.g., sensor noise, uncertainty in gust frequency, etc..

Christopher Forte↗
NASA NTRSDate not supplied

Gust Load Alleviation Control and Gust Estimation for a High Aspect Ratio Wing Wind Tunnel Model

This paper presents the gust load alleviation (GLA) study of the aspect ratio 13.5 Common Research Model (CRM) wind-tunnel model. This study details the design of the GLA controller in preparation for wind-tunnel testing in the Transonic Dynamics Tunnel at NASA Langley Research Center. An aeroservolastic (ASE) model is first reduced using a model reduction method that takes advantage of the sinusoidal steady-state response. Then, the reduced model is used to design an extended-state Kalman filter which estimates the states and the sinusoidal gust input. The GLA control is then derived using the optimal control solution to a multi-objective cost function. The results of the GLA controller indicate a 69.07% reduction in wing-root strain without sensor noise and 68.45% reduction with sensor noise, while maintaining robust stability margins. An adaptive GLA controller is developed for uncertain gust frequency and shows a 71.03% reduction in wing root strain compared to the non-adaptive control reduction of just 39.19%.

Christopher Forte↗
NASA NTRS2003

Contributions of the NASA Langley Research Center to the DARPA/AFRL/NASA/ Northrop Grumman Smart Wing Program

An overview of the contributions of the NASA Langley Research Center (LaRC) to the DARPA/AFRL/NASA/ Northrop Grumman Corporation (NGC) Smart Wing program is presented. The overall objective of the Smart Wing program was to develop smart** technologies and demonstrate near-flight-scale actuation systems to improve the aerodynamic performance of military aircraft. NASA LaRC s roles were to provide technical guidance, wind-tunnel testing time and support, and Computational Fluid Dynamics (CFD) analyses. The program was divided into two phases, with each phase having two wind-tunnel entries in the Langley Transonic Dynamics Tunnel (TDT). This paper focuses on the fourth and final wind-tunnel test: Phase 2, Test 2. During this test, a model based on the NGC Unmanned Combat Air Vehicle (UCAV) concept was tested at Mach numbers up to 0.8 and dynamic pressures up to 150 psf to determine the aerodynamic performance benefits that could be achieved using hingeless, smoothly-contoured control surfaces actuated with smart materials technologies. The UCAV-based model was a 30% geometric scale, full-span, sting-mounted model with the smart control surfaces on the starboard wing and conventional, hinged control surfaces on the port wing. Two LaRC-developed instrumentation systems were used during the test to externally measure the shapes of the smart control surface and quantify the effects of aerodynamic loading on the deflections: Videogrammetric Model Deformation (VMD) and Projection Moire Interferometry (PMI). VMD is an optical technique that uses single-camera photogrammetric tracking of discrete targets to determine deflections at specific points. PMI provides spatially continuous measurements of model deformation by computationally analyzing images of a grid projected onto the model surface. Both the VMD and PMI measurements served well to validate the use of on-board (internal) rotary potentiometers to measure the smart control surface deflection angles. Prior to the final entry, NASA LaRC also performed three-dimensional unstructured Navier Stokes CFD analyses in an attempt to predict the potential aerodynamic impact of the smart control surface on overall model forces and moments. Eight different control surface shapes were selected for study at Mach = 0.6, Reynolds number = 3.25 x 10(exp 6), and + 2 deg., 3 deg., 8 deg., and 10 deg.model angles-of-attack. For the baseline, undeflected control surface geometry, the CFD predictions and wind-tunnel results matched well. The agreement was not as good for the more complex aero-loaded control surface shapes, though, because of the inability to accurately predict those shapes. Despite these results, the NASA CFD study served as an important step in studying advanced control effectors.

Florance, Jennifer P.↗
NASA NTRS1992

Design, test, and evaluation of three active flutter suppression controllers

Three control law design techniques for flutter suppression are presented. Each technique uses multiple control surfaces and/or sensors. The first method uses traditional tools (such as pole/zero loci and Nyquist diagrams) for producing a controller that has minimal complexity and which is sufficiently robust to handle plant uncertainty. The second procedure uses linear combinations of several accelerometer signals and dynamic compensation to synthesize the model rate of the critical mode for feedback to the distributed control surfaces. The third technique starts with a minimum-energy linear quadratic Gaussian controller, iteratively modifies intensity matrices corresponding to input and output noise, and applies controller order reduction to achieve a low-order, robust controller. The resulting designs were implemented digitally and tested subsonically on the active flexible wing wind-tunnel model in the Langley Transonic Dynamics Tunnel. Only the traditional pole/zero loci design was sufficiently robust to errors in the nominal plant to successfully suppress flutter during the test. The traditional pole/zero loci design provided simultaneous suppression of symmetric and antisymmetric flutter with a 24-percent increase in attainable dynamic pressure. Posttest analyses are shown which illustrate the problems encountered with the other laws.

Adams, William M., Jr.↗
NASA NTRS1980

Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

An aspect ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumental with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Results from initial wind tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting surface theory.

Sandford, M. C.↗
NASA NTRS1980

Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

An aspect-ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumented with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Selected results from initial wind-tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing-edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting-surface theory.

Sandford, M. C.↗
NASA NTRS1992

Rolling maneuver load alleviation using active controls

Rolling Maneuver Load Alleviation (RMLA) was demonstrated on the Active Flexible Wing (AFW) wind tunnel model in the LaRC Transonic Dynamics Tunnel. The design objective was to develop a systematic approach for developing active control laws to alleviate wing incremental loads during roll maneuvers. Using linear load models for the AFW wind-tunnel model which were based on experimental measurements, two RMLA control laws were developed based on a single-degree-of-freedom roll model. The RMLA control laws utilized actuation of outboard control surface pairs to counteract incremental loads generated during rolling maneuvers and roll performance. To evaluate the RMLA control laws, roll maneuvers were performed in the wind tunnel at dynamic pressures of 150, 200, and 250 psf and Mach numbers of .33, .38, and .44, respectively. Loads obtained during these maneuvers were compared to baseline maneuver loads. For both RMLA controllers, the incremental torsion moments were reduced by up to 60 percent at all dynamic pressures and performance times. Results for bending moment load reductions during roll maneuvers varied. In addition, in a multiple function test, RMLA and flutter suppression system control laws were operated simultaneously during roll maneuvers at dynamic pressures 11 percent above the open-loop flutter dynamic pressure.

Woods-Vedeler, Jessica A.↗