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Tuned flaplike tab on trailing edge of helicopter rotor blade reduces blade vibration. Flapping tab on rotor blade produces aerodynamic loads that oppose vibrational aerodynamic loads. With proper choice of tab parameters, tab made to cancel much of blade vibration over wide range of rotor speeds and airspeeds.
A three-dimensional time dependent full potential aerodynamic analysis of single rotation propellers has been developed. The primary purpose of the code is to provide a capability of doing propfan aeroelastic analysis in the nonlinear transonic regime. A secondary purpose is to provide a validation of the unsteady lifting surface panel method that has been have developed. Results will be shown for steady state aerodynamic loading, unsteady aerodynamic response to forced aeroelastic deformations, and free aeroelastic response. Comparisons are made to experimental data and corresponding panel code results.
A finite element formulation is developed to analyze large-amplitude panel flutter of arbitrary laminated plates. The plates considered are anisotropic composite, thin rectangular panels. The equations of motion for an oscillating plate are determined and solved by linearizing the nonlinear stiffness matrices. The solution procedure is presented to determine the limit-cycle motions which are caused by the large deflections and vibrations induced by the areodynamic load. The aerodynamic load is defined by the first-order piston theory. Examples studied include cross-ply laminates with various numbers of layers and three-layer angle-ply laminates with different lamination angles. The effects of simply supported and clamped boundary conditions of a cross-ply laminate are also examined.
Flutter characteristics tested on cylindrical segments of longitudinal arrays of rectangular curved panels under compressive loading and aerodynamic heating
Wind load aerodynamics effects on launch vehicle flight control systems design
An objective of the HPCC Program at NASA Langley has been to promote the use of advanced computing techniques to more rapidly solve the problem of multidisciplinary optimization of a supersonic transport configuration. As a result, a software system has been designed and is being implemented to integrate a set of existing discipline analysis codes, some of them CPU-intensive, into a distributed computational framework for the design of a High Speed Civil Transport (HSCT) configuration. The proposed paper will describe the engineering aspects of integrating these analysis codes and additional interface codes into an automated design system. The objective of the design problem is to optimize the aircraft weight for given mission conditions, range, and payload requirements, subject to aerodynamic, structural, and performance constraints. The design variables include both thicknesses of structural elements and geometric parameters that define the external aircraft shape. An optimization model has been adopted that uses the multidisciplinary analysis results and the derivatives of the solution with respect to the design variables to formulate a linearized model that provides input to the CONMIN optimization code, which outputs new values for the design variables. The analysis process begins by deriving the updated geometries and grids from the baseline geometries and grids using the new values for the design variables. This free-form deformation approach provides internal FEM (finite element method) grids that are consistent with aerodynamic surface grids. The next step involves using the derived FEM and section properties in a weights process to calculate detailed weights and the center of gravity location for specified flight conditions. The weights process computes the as-built weight, weight distribution, and weight sensitivities for given aircraft configurations at various mass cases. Currently, two mass cases are considered: cruise and gross take-off weight (GTOW). Weights information is obtained from correlations of data from three sources: 1) as-built initial structural and non-structural weights from an existing database, 2) theoretical FEM structural weights and sensitivities from Genesis, and 3) empirical as-built weight increments, non-structural weights, and weight sensitivities from FLOPS. For the aeroelastic analysis, a variable-fidelity aerodynamic analysis has been adopted. This approach uses infrequent CPU-intensive non-linear CFD to calculate a non-linear correction relative to a linear aero calculation for the same aerodynamic surface at an angle of attack that results in the same configuration lift. For efficiency, this nonlinear correction is applied after each subsequent linear aero solution during the iterations between the aerodynamic and structural analyses. Convergence is achieved when the vehicle shape being used for the aerodynamic calculations is consistent with the structural deformations caused by the aerodynamic loads. To make the structural analyses more efficient, a linearized structural deformation model has been adopted, in which a single stiffness matrix can be used to solve for the deformations under all the load conditions. Using the converged aerodynamic loads, a final set of structural analyses are performed to determine the stress distributions and the buckling conditions for constraint calculation. Performance constraints are obtained by running FLOPS using drag polars that are computed using results from non-linear corrections to the linear aero code plus several codes to provide drag increments due to skin friction, wave drag, and other miscellaneous drag contributions. The status of the integration effort will be presented in the proposed paper, and results will be provided that illustrate the degree of accuracy in the linearizations that have been employed.
This work details the application of a lattice-Boltzmann method–very-large-eddy simulation (LBM-VLES) employed by the software suite, PowerFLOW. This LBM-VLES simulation predicted the aeroacoustic noise emanating from a representative, small unmanned aircraft system rotor, namely, the DJI-9450 in a hover condition. Predicted total aerodynamic loading as well as 2D aerodynamic loading along discrete spanwise sections of a rotor blade were compared to lower fidelity predictions and experimental results acquired in the Structural Acoustic Loads and Transmission anechoic chamber facility at the NASA Langley Research Center. The total acoustic spectra were decomposed into tonal and broadband components, which showed that broadband noise was a dominant contributor above 1 kHz for this rotor. These data were then compared to experimentally acquired data, showing good agreement up to approximately 11 kHz. Above 11 kHz, however, a grid sensitivity study showed dependency of the highest resolvable frequency on the spatial resolution of the computational domain, explaining the roll off in predicted data. Individual broadband noise sources were further investigated by calculating one-third octave sound pressure levels of the unsteady pressure fluctuations acting on the rotor, providing evidence that blade self-noise was the prominent noise source. Using these results, blade wake interaction noise was seen to be negligible for this particular rotor, which was further validated by calculating blade vortex miss distances and comparing to theory.
The origin of one/rev rotor aerodynamic loads which arise in tiltrotor aircraft during airplane-mode high speed pull-up and push-over maneuvers is examined using a coupled rotor/fuselage dynamic simulation. A modified eigenstructure assignment technique is used to design a controller which alleviates the in-plane loads during high pitch rate maneuvers. The controller utilizes rotor cyclic pitch inputs to restructure the aircraft short period and phugoid responses in order to achieve the coupling between pitch rate and rotor flapping responses which minimizes the rotor aerodynamic loading. Realistic time delays in the feedback path are considered during the controller design. Stability robustness in the presence of high frequency modeling errors is ensured through the use of singular value analysis.
A simplified procedure is shown for calculating the once-per-revolution oscillating aerodynamic thrust loads on propellers of tractor airplanes at zero yaw. The only flow field information required for the application of the procedure is a knowledge of the upflow angles at the horizontal center line of the propeller disk. Methods are presented whereby these angles may be computed without recourse to experimental survey of the flow field. The loads computed by the simplified procedure are compared with those computed by a more rigorous method and the procedure is applied to several airplane configurations which are believed typical of current designs. The results are generally satisfactory.
Stability & control of aircraft - aerodynamic loading
A static aeroelastic analysis capability that can predict aerodynamic loads for the deformed shape of the composite wing has been developed. The finite element method (FEM) was used for composite plate structural analysis, and the linear vortex lattice method (VLM) was used for steady aerodynamic analysis. The final deformed shape of the wing due to applied forces is determined by iterative manner using FEM and VLM. FEM and VLM analysis are related by a surface spline interpolation procedure. The wing with Gr/Ep composite material has been investigated to see the wing deformation effect. Aerodynamic load change due to wing flexibility has been investigated. Also, the effect of fiber orientation and sweep angle on the deformation pattern and aerodynamic coefficients are examined. For a certain fiber orientation, the deflection and aerodynamic loading of the composite wing is very much reduced. The swept forward wing has more significant effect of wing flexibility on aerodynamic coefficient than the swept back wing does.
A highly efficient numerical method is developed for three-dimensional, periodic, vortical flows around a cascade of loaded airfoils. The method linearizes Euler's equations about the mean flow of the cascade and thus fully accounts for the effects of distortion of the vortical disturbances as they propogate and interact with the cascade mean flow. The numerical scheme is based on splitting the unsteady velocity into vortical and potential parts. The latter is governed by a non-constant coefficient inhomogeneous convective wave equation. A new and computationally suitable out-flow conditions are derived and avoid the difficulties associated with the singular velocity downstream. Solutions were obtained in the frequency domain by using a body-fitted coordinate system. Results are presented to demonstrate the effects of the out-flow boundary conditions, cascade spacing, mean blade loading and gust upstream conditions on the aerodynamics response and unsteady pressure field of a cascade.
Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.
Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.
The aerodynamic load characteristics and the performance degradation of moderate aspect ratio wings and rotors with simulated glaze leading-edge ice have been studied using a three-dimensional, compressible Navier-Stokes solver. The effect of a splitter plate at the wing root on both clean and iced wing configurations has been studied and the results are compared with the experiment. A significant difference has been observed with and without splitter plates in the magnitude of flow separation and aerodynamic loading at the inboard stations for the iced wing at 8-deg angle of attack. Inviscid calculations were performed and compared with viscous calculations to investigate whether the performance of iced swept wings can be inexpensively predicted using Euler methods. It is shown that inviscid calculations predict higher aerodynamic loading than viscous calculations, and cannot model separation effects. A typical nonlifting helicopter rotor in forward flight condition is also studied, and the penalty due to the leading-edge ice formation on the required torque is numerically demonstrated.
At the most fundamental level, main rotor loading noise is caused by the harmonically-varying aerodynamic loads (acoustic pressures) exerted by the rotating blades on the air. Rotorcraft main rotor noise is therefore, in principle, a function of rotor control inputs, and thus the forces and moments required to achieve steady, or "trim", flight equilibrium. In certain flight conditions, the ensuing aerodynamic loading on the rotor(s) can result in highly obtrusive harmonic noise. The effect of the propulsive force, or X-force, on Blade-Vortex Interaction (BVI) noise is well documented. This paper presents an acoustics parametric sensitivity analysis of the effect of varying rotor aerodynamic pitch hub trim moments on BVI noise radiated by an S-70 helicopter main rotor. Results show that changing the hub pitching moment for an isolated rotor, trimmed in nominal 80 knot, 6 and 12 deg descent, flight conditions, alters the miss distance between the blades and the vortex in ways that have varied and noticeable effects on the BVI radiated-noise directionality. Peak BVI noise level is however not significantly altered. The application of hub pitching moment allows the attitude of the fuselage to be controlled; for example, to compensate for the uncomfortable change in fuselage pitch attitude introduced by a fuselage-mounted X-force controller.
Proprotor Aeroelastic Stability Analysis, now at version 4.5 (PASTA 4.5), is a FORTRAN computer program for analyzing the aeroelastic stability of a tiltrotor aircraft in the airplane mode of flight. The program employs a 10-degree- of-freedom (DOF), discrete-coordinate, linear mathematical model of a rotor with three or more blades and its drive system coupled to a 10-DOF modal model of an airframe. The user can select which DOFs are included in the analysis. Quasi-steady strip-theory aerodynamics is employed for the aerodynamic loads on the blades, a quasi-steady representation is employed for the aerodynamic loads acting on the vibrational modes of the airframe, and a stability-derivative approach is used for the aerodynamics associated with the rigid-body DOFs of the airframe. Blade parameters that vary with the blade collective pitch can be obtained by interpolation from a user-defined table. Stability is determined by examining the eigenvalues that are obtained by solving the coupled equations of motions as a matrix eigenvalue problem. Notwithstanding the relative simplicity of its mathematical foundation, PASTA 4.5 and its predecessors have played key roles in a number of engineering investigations over the years.
The general equations describing the nonlinear fluttering oscillations of shallow, curved, heated orthotropic panels have been derived. The formulation takes into account the location of the panel on the surface of a generic hypersonic vehicle, when calculating the aerodynamic loads. It is also shown that third order piston theory produces unsteady aerodynamic loading which is in close agreement with that based upon direct solution of the Euler equations. Results, for simply supported panels, are obtained using Galerkin's method combined with direct numerical integration in time to compute stable limit cycle amplitudes. These results illustrate the sensitivity of the aeroelastic behavior to the unsteady aerodynamic assumptions, temperature, orthotropicity and flow orientation.