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NASA NTRS1989

Aeroservoelastic wind-tunnel investigations using the active flexible wing model - Status and recent accomplishments

This paper describes the status of the joint NASA/Rockwell Active Flexible Wing Wind-Tunnel Test Program. The objectives of the program are to develop and validate the analysis, design and test methodologies required to apply multifunction active control technology for improving aircraft performance and stability. Major tasks of the program include designing digital multiinput/multioutput flutter-suppression and rolling-maneuver-load-alleviation concepts for a flexible full-span wind-tunnel model, obtaining an experimental data base for the basic model and each control concept, and providing comparisons between experimental and analytical results to validate the methodologies. This program is also providing the opportunity to improve real-time simulation techniques and to gain practical experience with digital control law implementation procedures.

Noll, Thomas↗
NASA NTRS1989

Aeroservoelastic modeling and applications using minimum-state approximations of the unsteady aerodynamics

The theoretical basis of flexible-aircraft modeling techniques encompassing aerodynamic, control, and elastic-structure effects is investigated analytically, with a focus on methods which employ minimum-state approximations for the unsteady aerodynamics. Rational-function approximations to generalized aerodynamic forces are reviewed; constraints and lag-coefficient optimization are explained; the problem of physical weighting in the minimum-state equations of motion is examined; and results of typical analyses from the NASA Active Flexible Wing project (Perry et al., 1988) are presented in extensive tables and graphs and discussed in detail. The minimum-state approach is shown to produce accurate models at significantly reduced computation costs.

Tiffany, Sherwood H.↗
NASA NTRS1989

Control law synthesis and optimization software for large order aeroservoelastic systems

A flexible aircraft or space structure with active control is typically modeled by a large-order state space system of equations in order to accurately represent the rigid and flexible body modes, unsteady aerodynamic forces, actuator dynamics and gust spectra. The control law of this multi-input/multi-output (MIMO) system is expected to satisfy multiple design requirements on the dynamic loads, responses, actuator deflection and rate limitations, as well as maintain certain stability margins, yet should be simple enough to be implemented on an onboard digital microprocessor. A software package for performing an analog or digital control law synthesis for such a system, using optimal control theory and constrained optimization techniques is described.

Mukhopadhyay, V.↗
NASA NTRS1990

Aeroservoelastic tailoring for lateral control enhancement

The application of an automated optimization approach to the integrated structural and control design of an aircraft wing and its control surface is illustrated. The performance index or cost function for optimization is the aileron hinge moment required to sustain a specified terminal roll rate at a chosen design speed at which aeroelastic effects are significant. The design variables used for aileron hinge-moment minimization are the aileron flap-to-chord ratio; the location of the aileron with respect to the roll axis; and the orientations of three advanced composite plies comprising 60 percent of the wing structure. The effects of aileron location and advanced composite ply orientation on the terminal roll rate and aileron hinge moment are examined. Optimization is seen to involve a compromise between a laminated structure/aileron combination that is effective in creating large rolling moments and one that minimizes the damping-in-roll. These two requirements are in conflict; the optimization driver must locate a compromise design point by moving the control surface and changing its size while also creating a suitable structural design.

Weisshaar, Terrence A.↗
NASA NTRS1991

Aeroservoelastic stabilization techniques for hypersonic flight vehicles

The potential of Hybrid Phase Stabilization (HPS), particularly for highly unstable aircraft, using a hypersonic flight vehicle (HSV) as a relevant example, is discussed. The development of HPS is presented and the result is compared with that generated using a conventional gain stabilization technique. Since HPS was not addressed in the MIL-spec requirements, a preliminary residual response metric was developed to provide guidance in assessing HPS.

Chan, Samuel Y.↗
NASA NTRS1991

Aeroservoelastic stabilization techniques for hypersonic flight vehicles

The potential of Hybrid Phase Stabilization (HPS), particularly for highly unstable aircraft, using a hypersonic flight vehicle (HSV) as a relevant example, is discussed. The development of HPS is presented and the result is compared with that generated using a conventional gain stabilization technique. Since HPS was not addressed in the MIL-spec requirements, a preliminary residual response metric was developed to provide guidance in assessing HPS.

Cheng, Peter Y.↗
NASA NTRS1992

Advanced aeroservoelastic stabilization techniques for hypersonic flight vehicles

Advanced high performance vehicles, including Single-Stage-To-Orbit (SSTO) hypersonic flight vehicles, that are statically unstable, require higher bandwidth flight control systems to compensate for the instability resulting in interactions between the flight control system, the engine/propulsion dynamics, and the low frequency structural modes. Military specifications, such as MIL-F-9490D and MIL-F-87242, tend to limit treatment of structural modes to conventional gain stabilization techniques. The conventional gain stabilization techniques, however, introduce low frequency effective time delays which can be troublesome from a flying qualities standpoint. These time delays can be alleviated by appropriate blending of gain and phase stabilization techniques (referred to as Hybrid Phase Stabilization or HPS) for the low frequency structural modes. The potential of using HPS for compensating structural mode interaction was previously explored. It was shown that effective time delay was significantly reduced with the use of HPS; however, the HPS design was seen to have greater residual response than a conventional gain stablized design. Additional work performed to advance and refine the HPS design procedure, to further develop residual response metrics as a basis for alternative structural stability specifications, and to develop strategies for validating HPS design and specification concepts in manned simulation is presented. Stabilization design sensitivity to structural uncertainties and aircraft-centered requirements are also assessed.

Chan, Samuel Y.↗
NASA NTRS1992

Aeroservoelastic stabilization technique refinement for hypersonic flight vehicles

Conventional gain-stabilization techniques introduce low frequency effective time delays which can be troublesome from the viewpoint of SSTOV vehicles' flying qualities. These time delays can be alleviated through a blending of gain-stabilization and phase-stabilization techniques; the resulting hybrid phase stabilization (HPS) for the low-frequency structural modes has been noted to have greater residual response than a conventional gain-stabilizer design. HPS design procedures are presently refined, and residual response metrics are developed.

Cheng, Peter Y.↗
NASA NTRS1993

ISAC - A tool for aeroservoelastic modeling and analysis

This paper discusses the capabilities of the Interaction of Structures, Aerodynamics, and Controls (ISAC) system of program modules. The major modeling, analysis, and data management components of ISAC are identified. Equations of motion are displayed for a Laplace-domain representation of the unsteady aerodynamic forces. Options for approximating a frequency-domain representation of unsteady aerodynamic forces with rational functions of the Laplace variable are shown. Linear time invariant state-space equations of motion that result are discussed. Model generation and analyses of stability and dynamic response characteristics are shown for an aeroelastic vehicle which illustrate some of the capabilities of ISAC as a modeling and analysis tool for aeroelastic applications.

Adams, William M., Jr.↗
NASA NTRS1990

Integrated aeroservoelastic synthesis for roll control

The objective of this study is to illustrate an integrated, parallel design procedure for optimal structural, aerodynamic, and aileron synthesis of an aircraft wing. The effects of combining weight minimization with structural tailoring (ply orientation and thickness) of a lifting surface, together with the wing geometry (sweep angle and taper ratio), and the aileron geometry (spanwise location and chordwise size) upon the lateral control effectiveness are discussed. Several optimization studies for the minimization of aileron hinge moment and wing weight, subject to a specified constant aircraft roll rate at a design airspeed (roll effectiveness), are performed.

Nam, Chang-Ho↗
NASA NTRS1993

ISAC: A tool for aeroservoelastic modeling and analysis

The capabilities of the Interaction of Structures, Aerodynamics, and Controls (ISAC) system of program modules is discussed. The major modeling, analysis, and data management components of ISAC are identified. Equations of motion are displayed for a Laplace-domain representation of the unsteady aerodynamic forces. Options for approximating a frequency-domain representation of unsteady aerodynamic forces with rational functions of the Laplace variable are shown. Linear time invariant state-space equations of motion that result are discussed. Model generation and analyses of stability and dynamic response characteristics are shown for an aeroelastic vehicle which illustrates some of the capabilities of ISAC as a modeling and analysis tool for aeroelastic applications.

Adams, William M., Jr.↗
NASA NTRS1997

Computational Aeroservoelastic Analysis with an Euler-Based Unsteady Flow Solver

The effectiveness of transpiration for simulation of structural deformations in steady and unsteady aeroelastic applications is examined. The majority of the investigations were performed using a highly integrated, finite-element code for the multidisciplinary analysis of flight vehicles. A supplement to this code, which allows for the generation of deflected meshes using modal superposition, was developed in this study. This research demonstrated that the transpiration boundary condition has strong potential for applications in unsteady aeroelastic analysis, such as in the prediction of flutter boundaries.

Arena, Andrew S., Jr.↗
NASA NTRS1999

Development and Validation of a Fast, Accurate and Cost-Effective Aeroservoelastic Method on Advanced Parallel Computing Systems

Progress to date towards the development and validation of a fast, accurate and cost-effective aeroelastic method for advanced parallel computing platforms such as the IBM SP2 and the SGI Origin 2000 is presented in this paper. The ENSAERO code, developed at the NASA-Ames Research Center has been selected for this effort. The code allows for the computation of aeroelastic responses by simultaneously integrating the Euler or Navier-Stokes equations and the modal structural equations of motion. To assess the computational performance and accuracy of the ENSAERO code, this paper reports the results of the Navier-Stokes simulations of the transonic flow over a flexible aeroelastic wing body configuration. In addition, a forced harmonic oscillation analysis in the frequency domain and an analysis in the time domain are done on a wing undergoing a rigid pitch and plunge motion. Finally, to demonstrate the ENSAERO flutter-analysis capability, aeroelastic Euler and Navier-Stokes computations on an L-1011 wind tunnel model including pylon, nacelle and empennage are underway. All computational solutions are compared with experimental data to assess the level of accuracy of ENSAERO. As the computations described above are performed, a meticulous log of computational performance in terms of wall clock time, execution speed, memory and disk storage is kept. Code scalability is also demonstrated by studying the impact of varying the number of processors on computational performance on the IBM SP2 and the Origin 2000 systems.

Goodwin, Sabine A.↗
NASA NTRS2003

Aeroservoelastic Model Validation and Test Data Analysis of the F/A-18 Active Aeroelastic Wing

Model validation and flight test data analysis require careful consideration of the effects of uncertainty, noise, and nonlinearity. Uncertainty prevails in the data analysis techniques and results in a composite model uncertainty from unmodeled dynamics, assumptions and mechanics of the estimation procedures, noise, and nonlinearity. A fundamental requirement for reliable and robust model development is an attempt to account for each of these sources of error, in particular, for model validation, robust stability prediction, and flight control system development. This paper is concerned with data processing procedures for uncertainty reduction in model validation for stability estimation and nonlinear identification. F/A-18 Active Aeroelastic Wing (AAW) aircraft data is used to demonstrate signal representation effects on uncertain model development, stability estimation, and nonlinear identification. Data is decomposed using adaptive orthonormal best-basis and wavelet-basis signal decompositions for signal denoising into linear and nonlinear identification algorithms. Nonlinear identification from a wavelet-based Volterra kernel procedure is used to extract nonlinear dynamics from aeroelastic responses, and to assist model development and uncertainty reduction for model validation and stability prediction by removing a class of nonlinearity from the uncertainty.

Brenner, Martin J.↗
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 NTRS2007

F15B-Quiet Spike Aeroservoelastic Flight Test Data Analysis

Airframe structural morphing technologies designed to mitigate sonic boom strength are being developed by Gulfstream Aerospace Corporation (GAC). Among these technologies is a concept in which an aircraft's frontend would be extended prior to supersonic acceleration. This morphing would effectively lengthen the vehicle, reducing peak sonic boom amplitude, but is also expected to partition the otherwise strong bow shock into a series of reduced-strength, non-coalescing shocklets. This combination of boom shaping techniques is predicted to transform the classic, high-impulse N-wave pattern typically generated by an aircraft traveling at supersonic speed into a signature more closely resembling a sinusoidal wave with a greatly reduced perceived loudness. 'QuietSpike' is GAC's nomenclature for its recently patented front-end vehicle morphing arrangement. The ability of Quiet Spike to effectively shape a vehicle's far- field sonic boom signature is highly dependent on the area distribution characteristics of the aircraft. The full aeroacoustic benefits of front-end morphing at farfield are only possible when the QuietSpike article and vehicle configuration are designed in consideration of each other. Adding QuietSpike technology to the airframe of an existing, non-boom-optimized supersonic vehicle is unlikely to result in an improved far-field signature due to the generally over-powering influence of wing- and inlet-generated shocks. Therefore, it is generally recognized within NASA and the industry that a clean-sheet vehicle design is required to demonstrate the theoretically predicted far-field aeroacoustic benefits of QuietSpike type morphing and other boom- mitigating concepts. NASA's Aeronautics Research Mission Directorate (ARMD) Supersonics Division has placed increased priority on near-term development and flight-testing of such a vehicle. To help achieve this objective, static and dynamic aerostructural proof-of-concept testing was considered a prudent step prior to a clean-sheet effort in order to reduce risk associated with a follow-on test program. Following a survey of potential test platforms, NASA Dryden's F-15B was selected as the target test vehicle primarily because of its unique ability to carry a largescale test apparatus to relevant supersonic flight speeds, so called the F15 -QS. The QuietSpike test article was constructed primarily of composite materials and attached to the forward fuselage of the F-1 5B bulkhead (see Figures 1,2). The QuietSpike test article replaces the current flight test noseboom and radome assembly. Power is supplied to the Quiet Spike motor assembly in order to extend and retract the Spike, and the Quiet Spike test article was appropriately instrumented with accelerometers, strain gages, pressure transducers, and thermocouples.

Brenner, Martin J.↗
NASA NTRS2015

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine, NASA Advanced Air Vehicles Program - Commercial Supersonic Technology Project - AeroServoElasticity

This paper covers the development of stage-by-stage and parallel flow path compressor modeling approaches for a Variable Cycle Engine. The stage-by-stage compressor modeling approach is an extension of a technique for lumped volume dynamics and performance characteristic modeling. It was developed to improve the accuracy of axial compressor dynamics over lumped volume dynamics modeling. The stage-by-stage compressor model presented here is formulated into a parallel flow path model that includes both axial and rotational dynamics. This is done to enable the study of compressor and propulsion system dynamic performance under flow distortion conditions. The approaches utilized here are generic and should be applicable for the modeling of any axial flow compressor design accurate time domain simulations. The objective of this work is as follows. Given the parameters describing the conditions of atmospheric disturbances, and utilizing the derived formulations, directly compute the transfer function poles and zeros describing these disturbances for acoustic velocity, temperature, pressure, and density. Time domain simulations of representative atmospheric turbulence can then be developed by utilizing these computed transfer functions together with the disturbance frequencies of interest.

Compressor Modeling Modeling↗