A method for determining aerodynamic damping of the structural bending modes with application to the Apollo/Saturn 1 launch vehicle
Comparison of damping derivatives and wind tunnel data of structural bending modes of Saturn 1
Meadows, M. T.↗
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Acoustic Mode Decomposition in Rectangular Ducts with Sheared Flow
The performance of new acoustic liner concepts are, in general, characterized and assessed in grazing flow rigs early in the development cycle. These test rigs expose an acoustic liner sample, installed on the side wall of the duct, to a grazing flow and incident acoustic field. The process to characterize these liners involves educing the impedance on the wall where the sample is installed and examining the acoustic power attenuation. Standard approaches to computing impedance or power attenuation generally consider only the effects of a 2D shear flow or uniform flow, on the acoustic field. In this study, the objective is to incorporate 3D shear flow effects in the analysis of acoustic mode attenuation in a rectangular duct flow rig. A modal analysis of microphone measurements obtained on the side walls of the duct upstream and downstream of the test section of the rig is developed. A Galerkin projection of the Pridmore-Brown equation is performed with Chebyshev basis functions in order to incorporate the effects of the Mach number profile on the computation of the axial wavenumber of each mode. Measurements of the Mach number profile are obtained in the test rig and used as input to compute the modes. Comparisons made between the sound field computed with traditional convective Helmholtz modes and the new procedure using Pridmore-Brown modes indicate that the computed acoustic field using Pridmore-Brown modes more accurately reconstructs the acoustic signal at each microphone in the array. The mode structure of the lowest-order mode is shown to be significantly impacted by shear flow refraction effects, and higher-order mode structures are also affected at higher frequencies and centerline Mach number. An assessment of the acoustic mode attenuation for two acoustic liner samples demonstrates that the computed mode amplitudes for both the traditional and new approach are in agreement for the lowest-order mode, but discrepancies arise when higher-order modes are the dominant component of the acoustic field.
Duct acoustics↗
Buffet induced structural/flight-control system interaction of the X-29A aircraft
High angle-of-attack flight regime research is currently being conducted for modern fighter aircraft at the NASA Ames Research Center's Dryden Flight Research Facility. This flight regime provides enhanced maneuverability to fighter pilots in combat situations. Flight research data are being acquired to compare and validate advanced computational fluid dynamic solutions and wind-tunnel models. High angle-of-attack flight creates unique aerodynamic phenomena including wing rock and buffet on the airframe. These phenomena increase the level of excitation of the structural modes, especially on the vertical and horizontal stabilizers. With high gain digital flight-control systems, this structural response may result in an aeroservoelastic interaction. A structural interaction on the X-29A aircraft was observed during high angle-of-attack flight testing. The roll and yaw rate gyros sensed the aircraft's structural modes at 11, 13, and 16 Hz. The rate gyro output signals were then amplified through the flight-control laws and sent as commands to the flaperons and rudder. The flight data indicated that as the angle of attack increased, the amplitude of the buffet on the vertical stabilizer increased, which resulted in more excitation to the structural modes. The flight-control system sensors and command signals showed this increase in modal power at the structural frequencies up to a 30 degree angle-of-attack. Beyond a 30 degree angle-of-attack, the vertical stabilizer response, the feedback sensor amplitude, and control surface command signal amplitude remained relatively constant. Data are presented that show the increased modal power in the aircraft structural accelerometers, the feedback sensors, and the command signals as a function of angle of attack. This structural interaction is traced from the aerodynamic buffet to the flight-control surfaces.
Voracek, David F.↗
The preferred-mode coherent structure in the near field of an axisymmetric jet with and without excitation
Visualization and phase-locked hot-wire measurements have been used to investigate the preferred-mode coherent structure in the near field of an axisymmetric jet for different Reynolds numbers and for the two limiting asymptotic states of the exit boundary layer: laminar and fully developed turbulent. Distributions of phase-averaged longitudinal and lateral velocities, coherent vorticity, coherent and incoherent turbulence Reynolds stresses, and coherent strain rate and production have been obtained over the spatial extent of the coherent structure for a Strouhal number of 0.3. It is found that these properties are independent of the initial conditions, but weakly dependent on the Reynolds number up to about 100,000. The naturally occurring coherent structure in the unexcited jet, educed via an optimized conditional sampling scheme, shows no clear dependence on the jet Reynolds number up to 800,000.
Hussain, A. K. M. F.↗
Recent Enhancements to the Development of CFD-Based Aeroelastic Reduced-Order Models
Recent enhancements to the development of CFD-based unsteady aerodynamic and aeroelastic reduced-order models (ROMs) are presented. These enhancements include the simultaneous application of structural modes as CFD input, static aeroelastic analysis using a ROM, and matched-point solutions using a ROM. The simultaneous application of structural modes as CFD input enables the computation of the unsteady aerodynamic state-space matrices with a single CFD execution, independent of the number of structural modes. The responses obtained from a simultaneous excitation of the CFD-based unsteady aerodynamic system are processed using system identification techniques in order to generate an unsteady aerodynamic state-space ROM. Once the unsteady aerodynamic state-space ROM is generated, a method for computing the static aeroelastic response using this unsteady aerodynamic ROM and a state-space model of the structure, is presented. Finally, a method is presented that enables the computation of matchedpoint solutions using a single ROM that is applicable over a range of dynamic pressures and velocities for a given Mach number. These enhancements represent a significant advancement of unsteady aerodynamic and aeroelastic ROM technology.
Silva, Walter A.↗
Analysis of nonlinear structures via mode synthesis
An effective procedure for NASTRAN was developed that permits any number of substructures of any size to be synthesized for the purpose of developing normal modes of vibration of the complete structural system. The technique is extended to permit modal transient analysis of the subdivided system. This latter procedure permits the use of NASTRAN's ability to include nonlinear forces in the problem. The five-phase process is accomplished using standard NASTRAN rigid formats with problem-independent alter packages and DMAP sequences.
Gieseke, R. K.↗
Comparison of candidate methods to distinguish noise modes from system modes in structural identification
In modal identification, nonphysical noise or computation modes always appear to help match the input-output data. This paper studies the ability of four criteria to distinguish which modes in a model are noise modes: (1) modal amplitude coherency, (2) the relative contribution of each mode to the pulse response indicated by the mode singular value, (3) the variances of the mode frequencies and damping factors produced by a chosen measurement noise level, and (4) identification of the backward-time in order to let the shift from positive to negative damping of the true system modes distinguish these modes from noise modes. Both simulated and experimental data are used to study the four criteria.
Longman, Richard W.↗
Expansion and orthogonalization of measured modes for structure identification
The purpose was to investigate a new simultaneous expansion/orthogonalization method in comparison with two previously published expansion methods and a widely used orthogonalization technique. Each expansion method uses data from an analytical model of the structure to complete the estimate of the mode shape vectors. Berman and Nagy used Guyan expansion in their work with improving analytical models. In this method, modes are expanded one at a time, producing a set not orthogonal with respect to the mass matrix. Baruch and Bar Itzhack's optimal orthogonalization procedure was used to subsequently adjust the expanded modes. A second expansion technique was presented by O'Callahan, Avitabile, and Reimer and separately by Kammer. Again, modes are expanded individually and orthogonalized after expansion with the same optimal technique as above. Finally, a simultaneous expansion/orthogonalization method was developed from the orthogonal Procrustes problem of computational mathematics. In this method modes are optimally expanded as a set and orthogonal with respect to the mass matrix as a result. Two demonstation problems were selected for the comparison of the methods described. The first problem is an 8 degree of freedom spring-mass problem first presented by Kabe. Several conditions were examined for expansion method including the presence of errors in the measured data and in the analysis models. As a second demonstration problem, data from tests of laboratory scale model truss structures was expanded for system identification. Tests with a complete structure produced a correlated analysis model and the stiffness and mass matrices. Tests of various damaged configurations produced measured data for 6 modes at 14 dof locations.
Smith, Suzanne Weaver↗
Simultaneous Excitation of Multiple-Input Multiple-Output CFD-Based Unsteady Aerodynamic Systems
A significant improvement to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) is presented. This improvement involves the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system that enables the computation of the unsteady aerodynamic state-space model using a single CFD execution, independent of the number of structural modes. Four different types of inputs are presented that can be used for the simultaneous excitation of the structural modes. Results are presented for a flexible, supersonic semi-span configuration using the CFL3Dv6.4 code.
Silva, Walter A.↗
Simultaneous Excitation of Multiple-Input Multiple-Output CFD-Based Unsteady Aerodynamic Systems
A significant improvement to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) is presented. This improvement involves the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system that enables the computation of the unsteady aerodynamic state-space model using a single CFD execution, independent of the number of structural modes. Four different types of inputs are presented that can be used for the simultaneous excitation of the structural modes. Results are presented for a flexible, supersonic semi-span configuration using the CFL3Dv6.4 code.
Silva, Walter A.↗
Adaptive Control of Flexible Structures Using Residual Mode Filters
Flexible structures containing a large number of modes can benefit from adaptive control techniques which are well suited to applications that have unknown modeling parameters and poorly known operating conditions. In this paper, we focus on a direct adaptive control approach that has been extended to handle adaptive rejection of persistent disturbances. We extend our adaptive control theory to accommodate troublesome modal subsystems of a plant that might inhibit the adaptive controller. In some cases the plant does not satisfy the requirements of Almost Strict Positive Realness. Instead, there maybe be a modal subsystem that inhibits this property. This section will present new results for our adaptive control theory. We will modify the adaptive controller with a Residual Mode Filter (RMF) to compensate for the troublesome modal subsystem, or the Q modes. Here we present the theory for adaptive controllers modified by RMFs, with attention to the issue of disturbances propagating through the Q modes. We apply the theoretical results to a flexible structure example to illustrate the behavior with and without the residual mode filter. We have proposed a modified adaptive controller with a residual mode filter. The RMF is used to accommodate troublesome modes in the system that might otherwise inhibit the adaptive controller, in particular the ASPR condition. This new theory accounts for leakage of the disturbance term into the Q modes. A simple three-mode example shows that the RMF can restore stability to an otherwise unstable adaptively controlled system. This is done without modifying the adaptive controller design.
Balas, Mark J.↗
Reduced-Order Aerodynamic Modeling Based on CFD Frequency Responses from Multisine Inputs
A system identification analysis was performed to determine a reduced-order model (ROM) of a computational fluid dynamics (CFD) solver in support of linear aeroservoelastic model development and feedback control design. The approach was applied to the FUN3D code for the half-span wind tunnel test article used in the NASA-Boeing collaboration called the Integrated Adaptive Wing Technology Maturation (IAWTM) project. In a transonic flow condition, multiple inputs (11 structural mode displacements and 3 control surface deflections) were simultaneously excited with orthogonal phase-optimized multisines while multiple outputs (the corresponding 14 generalized aerodynamic forces) were recorded. From these recorded times series, the matrix of frequency responses was computed and subsequently fit using rational function approximations (RFAs). It was found that the entire (14 x 14) matrix of frequency responses could be determined from a single CFD run and that results generally followed trends predicted using other methods. Differences were attributed to the modeling fidelity and nonlinearities from structural mode and control surface interactions at higher reduced frequencies. More accurate fits of the RFAs to the frequency response data were obtained by making two CFD runs, one with only structural mode excitations and one with only control surface excitations, which reduced the degree of nonlinearity in the modeling data.
Aeroservoelasticity↗
Basic features of the STS/Spacelab vibration environment
The Space Shuttle acceleration environment is characterized. The acceleration environment is composed of a residual or quasi-steady component and higher frequency components induced by vehicle structural modes and the operation of onboard machinery. Quasi-steady accelerations are generally due to atmospheric drag, gravity gradient effects, and rotational forces. These accelerations tend to vary with the orbital frequency (approx. 10(exp -4) Hz) and have magnitudes less than or equal to 10(exp -6) g(sub 0) (where 1 g(sub 0) is terrestrial gravity). Higher frequency g-jitter is characterized by oscillatory disturbances in the 1-100 Hz range and transient components. Oscillatory accelerations are related to the response of large flexible structures like antennae, the Spacelab module, and the Orbiter itself, and to the operation of rotating machinery. The Orbiter structural modes in the 1-10 Hz range, are excited by oscillatory and transient disturbances and tend to dominate the energy spectrum of the acceleration environment. A comparison of the acceleration measurements from different Space Shuttle missions reveals the characteristic signature of the structural modes of the Orbiter overlaid with mission specific hardware induced disturbances and their harmonics. Transient accelerations are usually attributed to crew activity and Orbiter thruster operations. During crew sleep periods, the acceleration levels are typically on the order of 10(exp -6) g(sub 0) (1 micro-g). Crew work and exercise tend to raise the accelerations to the 10(exp -3) g(sub 0) (1 milli-g) level. Vernier reaction control system firings tend to cause accelerations of 10(exp -4) g(sub 0), while primary reaction control system and Orbiter maneuvering system firings cause accelerations as large as 10(exp -2) g(sub 0). Vibration isolation techniques (both active and passive systems) used during crew exercise have been shown to significantly reduce the acceleration magnitudes.
Baugher, Charles R.↗
Band-limited power flow into enclosures
Equations for the band-limited power flow to a cavity in the low-frequency regime are derived. The total power to cavity is obtained by summing separate calculations of the power from structural modes resonant in the band and power from structural modes resonant below the band. High-frequency relations compatible with the usual statistical energy analysis and generalized to include other excitations in addition to diffuse fields are also provided.
Pope, L. D.↗
GOES-R Active Vibration Damping Controller Design, Implementation, and On-Orbit Performance
GOES-R series spacecraft feature a number of flexible appendages with modal frequencies below 3.0 Hz which, if excited by spacecraft disturbances, can be sources of undesirable jitter perturbing spacecraft pointing. In order to meet GOES-R pointing stability requirements, the spacecraft flight software implements an Active Vibration Damping (AVD) rate control law which acts in parallel with the nadir point attitude control law. The AVD controller commands spacecraft reaction wheel actuators based upon Inertial Measurement Unit (IMU) inputs to provide additional damping for spacecraft structural modes below 3.0 Hz which vary with solar wing angle. A GOES-R spacecraft dynamics and attitude control system identified model is constructed from pseudo-random reaction wheel torque commands and IMU angular rate response measurements occurring over a single orbit during spacecraft post-deployment activities. The identified Fourier model is computed on the ground, uplinked to the spacecraft flight computer, and the AVD controller filter coefficients are periodically computed on-board from the Fourier model. Consequently, the AVD controller formulation is based not upon pre-launch simulation model estimates but upon on-orbit nadir point attitude control and time-varying spacecraft dynamics. GOES-R high-fidelity time domain simulation results herein demonstrate the accuracy of the AVD identified Fourier model relative to the pre-launch spacecraft dynamics and control truth model. The AVD controller on-board the GOES-16 spacecraft achieves more than a ten-fold increase in structural mode damping of the fundamental solar wing mode while maintaining controller stability margins and ensuring that the nadir point attitude control bandwidth does not fall below 0.02 Hz. On-orbit GOES-16 spacecraft appendage modal frequencies and damping ratios are quantified based upon the AVD system identification, and the increase in modal damping provided by the AVD controller for each structural mode is presented. The GOES-16 spacecraft AVD controller frequency domain stability margins and nadir point attitude control bandwidth are presented along with on-orbit time domain disturbance response performance.
Active Vibration Damping↗
Ground/Flight Correlation of Aerodynamic Loads with Structural Response
United States Air Force Research Laboratory (AFRL) ground tests at the NASA Transonic Dynamics Tunnel (TDT) and NASA flight tests provide a basis and methodology for in-flight characterization of the aeroelastic performance through the monitoring of the fluid-structure interaction using surface flow sensors. NASA NF-15B flight tests provided a unique opportunity to test the correlation of aerodynamic loads with sectional flow attachment/detachment points, also known as flow bifurcation points (FBPs), as observed in previous wind tunnel tests. The NF-15B tail was instrumented with hot-film sensors and strain gages for measuring root-bending strains. These data were gathered via selected sideslip maneuvers performed at level flight and subsonic speeds. The aerodynamic loads generated by the sideslip maneuver resulted in root-bending strains and hot-film sensor signals near the stagnation region that were highly correlated. For the TDT tests, a flexible wing section developed under the AFRL SensorCraft program was instrumented with strain gages, accelerometers, and hot-film sensors at multiple span stations. The TDT tests provided data showing a gradual phase change between the FBP and the structural mode occurred during a resonant condition as the wings structural modes were excited by the tunnel-generated gusts.
Mangalam, Arun S.↗
Simultaneous expansion and orthogonalization of measured modes for structure identification
Tests of large structures on-orbit will be performed with measurements at a relatively few structure points. Values for the unmeasured degrees of freedom (dofs) can be estimated based on measured dofs and analytical model dynamic information. These 'expanded' mode shapes are useful for optimal-update identification and damage location as well as test/analysis correlation. A new method of expansion for test mode shape vectors is developed from the orthogonal Procrustes problem from computational linear algebra. A subspace defined by the set of measured dofs is compared to a subspace defined by mode shapes from an analytical model of the structure. The method simultaneously expands and orthogonalizes the mode shape vectors. Two demonstration problems are used to compare the new method to current expansion techniques. One demonstration uses test data from a laboratory scale-model truss structure. Performance of the new method is comparable or superior to that of the previous expansion methods which require separate orthogonalization.
Smith, Suzanne Weaver↗
Ares-I-X Stability and Control Flight Test: Analysis and Plans
The flight test of the Ares I-X vehicle provides a unique opportunity to reduce risk of the design of the Ares I vehicle and test out design, math modeling, and analysis methods. One of the key features of the Ares I design is the significant static aerodynamic instability coupled with the relatively flexible vehicle - potentially resulting in a challenging controls problem to provide adequate flight path performance while also providing adequate structural mode damping and preventing adverse control coupling to the flexible structural modes. Another challenge is to obtain enough data from the single flight to be able to conduct analysis showing the effectiveness of the controls solutions and have data to inform design decisions for Ares I. This paper will outline the modeling approaches and control system design to conduct this flight test, and also the system identification techniques developed to extract key information such as control system performance (gain/phase margins, for example), structural dynamics responses, and aerodynamic model estimations.
Brandon, Jay M.↗