SEARCH · Engineering Papers

Results for “blade representation”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 73 records · Page 4

All sources NASA NTRS DOE OSTI

NASA NTRS1990

Shaft flexibility effects on the forced response of a bladed-disk assembly

A model analysis approach is used to study the forced response of an actual flexible bladed-disk-shaft system. Both in-plane and out-of-plane flexible deformations of the bladed-disk assembly are considered, in addition to its rigid-body translations and rotations, resulting from the bending of the supporting flexible shaft in two orthogonal planes. The effects of Coriolis forces and structural coupling between flexible and rigid disk motions on the system's response are investigated. Aerodynamic loads acting on the rotating and vibrating bladed-disk assembly are accounted for through a simple quasi-steady representation, to evaluate their influence, combined with shaft flexibility and Coriolis effects.

Khader, N.↗

NASA NTRS1994

Sensitivity of lag-damping correlations to structural and aerodynamic approximations of isolated experimental rotors in forward flight

The predictions of regressive lag-mode damping levels are correlated with the database of an isolated, soft-inplane, three-blade rotor operated untrimmed. The database was generated at the Army Aeroflightdynamics Directorate at Ames. The correlation covers a broad range of data, from near-zero thrust conditions in hover to high-thrust and highly stalled conditions in forward flight with advance ratio as high as 0.55 and shaft angle as high as 20 degrees. In the experimental rotor, the airfoil or blade portion has essentially uniform mass and stiffness distributions, but the root flexure has highly nonuniform mass and stiffness distributions. Accordingly, the structural approximations refer to four models of root-flexure-blade assembly. They range from a rigid flap-lag model to three elastic flap-lag-torsion models, which differ in modeling the root flexure. The three models of root-flexure are: three root springs in which the bending-torsion couplings are fully accounted for; a finite-length beam element with some average mass and stiffness distributions such that the fundamental frequencies match those of the experimental model; and accurate modal representation in which the actual mass and stiffness distributions of the experimental root-flexure-blade assembly are used in calculating the nonrotating mode shapes. The four models of root-flexure-blade assembly are referred to as the rigid flap-lag model, spring model, modified model and modal model. For each of these four models of the root-flexure-blade assembly, the predictions are based on the following five aerodynamic theories: ear theory, which accounts for large angle-of-attack and reverse-flow effects on lift, and has constant drag and pitching moment; quasisteady stall theory, which includes quasisteady stall lift, drag and pitching moment characteristics of the airfoil section, dynamics stall theory, which uses the ONERA dynamic stall models of lift, drag and pitching moment; dynamic wake theory, which is based on a finite-state three-dimensional wake model and includes all wake effects including both shed and trailing vorticity; and dynamics and wake theory, which combines both dynamic stall theory and dynamic wake theory and is a relatively complete aerodynamic representation.

Gaonkar, G. H.↗

NASA NTRS1974

Computer experiments on periodic systems identification using rotor blade transient flapping-torsion responses at high advance ratio

Systems identification methods have recently been applied to rotorcraft to estimate stability derivatives from transient flight control response data. While these applications assumed a linear constant coefficient representation of the rotorcraft, the computer experiments described in this paper used transient responses in flap-bending and torsion of a rotor blade at high advance ratio which is a rapidly time varying periodic system.

Hohenemser, K. H.↗

NASA NTRS1985

Three-dimensional tomographic reconstruction of the flow around a revolving helicopter rotorblade - A numerical simulation

This paper discusses tomographic data acquisition and reconstruction procedures for determining the flow around a revolving helicopter rotor blade. The data acquisition procedure is simulated using, as a testbed, a numerical representation of the flow and projections through this known field are computed. An investigation has been made of: (1) the effect of the number of projections on the accuracy of the reconstruction and (2) the fidelity of the reconstruction when the region of intersection of all projections is smaller than the spatial extent of the flow.

Hesselink, L.↗

NASA NTRS1992

Helicopter response to atmospheric turbulence

A new time-domain method for simulating cyclostationary turbulence as seen by a translating and rotating blade element has recently been developed for the case of one-dimensional spectral distribution. This paper extends the simulation method to the cases of two- and three-dimensional spectral distributions and presents validation results for the two-dimensional case. The statistics of an isolated rigid blade flapping response to turbulence are computed using a two-dimensional spectral representation of the von Karman turbulence model, and the results are compared with those obtained using the conventional space-fixed turbulence analysis. The new turbulence simulation method is used for predicting the Black Hawk helicopter response to atmospheric turbulence.

Riaz, J.↗

DOE OSTI2023

Capturing Time-Varying Wake Dynamics Using Hybridized Actuator Disks in Steady-State Simulations for Improved Optimization Efficiency

When optimizing turbine blade properties, using a time varying, unsteady simulation allows for high-fidelity representations of the wake dynamics. Unfortunately these types of optimizations are prohibitively expensive in terms of computation time and memory requirements. This presentation aims to show a method of transferring the wake dynamics of an unsteady simulation to a much more computationally tractable steady state simulation using hybrid actuator disks informed by unsteady, actuator line model dynamics.

actuator disk model↗

DOE OSTI2024

A Solution Method for the Filtered Lifting Line Theory

The filtered lifting line theory presents a continuous form of the inviscid momentum equations of flow over a lifting device, such as a wing or rotor blade, using body forces without mathematical singularities. This theory is also consistent with an actuator line representation of a lifting device. In this work, we present a reformulation of the equations in terms of the local flow angle along the line, which allows solving the stand-alone equations using multivariate root-finding algorithms. This approach can be used to obtain a fast, computationally inexpensive solution of the loading distribution along a wing without the need to perform computational fluid dynamic simulations. We study the requirements in terms of resolution in the spanwise direction and establish the criteria for spacing and minimum amount of points required along the blade to obtain converged solutions. The solutions are compared to results from large-eddy simulations, and we observed excellent agreement with less than a percent difference in quantities along the blade between the methods.

17 WIND ENERGY↗

NASA NTRS1984

The 3-D inelastic analysis methods for hot section components (base program)

A series of computer codes are sought that would permit more accurate and efficient three-dimensional inelastic structural analysis of combustor liners, turbine blades, and turbine vanes. Each code embodies a progression of mathematical models for increasingly comprehensive representation of the geometrical features, loading conditions, and forms of nonlinear material response that distinguish these three groups of hot section components.

Todd, E. S.↗

NASA NTRS1987

All-theoretical prediction of cabin noise due to impingement of propeller vortices on a wing structure

Reported calculations of structure-borne cabin noise for a small twin engine aircraft powered by tractor propellers rely on the following three-stage methodological breakup of the problem: (1) the unsteady-aerodynamic prediction of wing lift harmonics caused by the whipping action of the vortex system trailed from each propeller; (2) the associated wing/fuselage structural response; (3) the cabin noise field for the computed wall vibration. The first part--the estimate of airloads--skirts a full-fledged aeroelastic situation by assuming the wing to be fixed in space while cancelling the downwash field of the cutting vortices. The model is based on an approximate high-frequency lifting-surface theory justified by the blade rate and flight Mach number of application. Its results drive a finite-element representation of the wing accounting for upper and lower skin surfaces, spars, ribs, and the presence of fuel. The fuselage, modeled as a frame-stiffened cylindrical shell, is bolted to the wing.

Martinez, R.↗

NASA NTRS1986

On 3-D inelastic analysis methods for hot section components

The objective of this program is to produce a series of new computer codes that permit more accurate and efficient three-dimensional inelastic structural analysis of combustor liners, turbine blades, and turbine vanes. Each code embodies a progression of mathematical models for increasingly comprehensive representation of the geometrical features, loading conditions, and forms of nonlinear material response that distinguish these three groups of hot section components.

Todd, E. S.↗

NASA NTRS1987

Three-dimensional inelastic analysis methods for hot section components

Todd, E. S.↗

NASA NTRS1987

Three dimensional flows in multistage compressors

Current multi-stage turbomachinery design/analysis methods are based on a time-averaged, axisymmetric representation of the flow field. The actual flow field is asymmetric and unsteady due to blade row interactions. The Reynolds averaged Navier-Stokes solvers are limited to single-stage machines for existing computers. Therefore, advanced multi-stage compressors will operate far off-design for portions of the flight regime. The objectives are to provide an experimentally validated average-passage calculation of multistage compressor blade row interactions and an experimentally validated time-accurate calculation of multi-stage compressor blade row interactions. Various aspects of this investigation are presented in viewgraph form.

Hathaway, Michael D.↗

NASA NTRS1983

Helicopter rotor loads using discretized matched asymptotic expansions

The numerical practicality of a matched asymptotic expansion approach for the computation of unsteady three dimensional airloads on a helicopter rotor was improved. This effort utilizes a discretized repesentation of the doublet strength distribution and helical streamlines to decrease the computational requirements of the original analysis. The continuous variation of the doublet strength was approximated by piecewise constant or piecewise quadratic distributions, and the helical trajectory of a fluid particle was approximated by connected straight line segments. As a direct result of these simplified representations the computational time required for the execution of a typical flight condition was reduced by an order of magnitude with respect to the requirements of the original analysis. Airloads which were computed using the discretized method for a two bladed model rotor and a full scale four bladed rotor are in close agreement with measured results and airloads from the original asymptotic analysis. For conditions characterized by significant rotor/wake interaction the piecewise constant representation requires a reduced azimuth spacing to maintain acceptable accuracy.

Pierce, G. A.↗

NASA NTRS1985

Unsteady analysis of rotor blade tip flow

The development of the VSAERO-TS and VSAERO-H computer programs for calculating the unsteady aerodynamic characteristics of arbitrarily shaped wings oscillating in pitch is presented. The effect of several wake parameters on chordwise pressure distribution in VSAERO-TS is given and the convergence characteristics of both programs are discussed. In the program, the influence coefficient for each panel is formulated for a planar surface and so a skewed panel is represented by a projected flat quadrilateral lying in the mean plane. Since panels in the extreme roll-up region of the tip vortex are highly skewed the program was modified to treat each highly skewed panel as a pair of triangles. The programs are validated by comparing the chordwise pressure distribution of several blade tip planforms with experimental data. The comparison, for the most part, is good. The triangular panel representation improved the chordwise pressure distribution near the tip region for higher mean angle of attack.

Maskey, B.↗

NASA NTRS2004

Improving Rotor-Stator Interaction Noise Code Through Analysis of Input Parameters

There are two major sources of aircraft noise. The first is from the airframe and the second is from the engines. The focus of the acoustics branch at NASA Glenn is on the engine noise sources. There are two major sources of engine noise; fan noise and jet noise. Fan noise, produced by rotating machinery of the engine, consists of both tonal noise, which occurs at discrete frequencies, and broadband noise, which occurs across a wide range of frequencies. The focus of my assignment is on the broadband noise generated by the interaction of fan flow turbulence and the stator blades. such as the sweep and stagger angles and blade count, as well as the flow parameters such as intensity of turbulence in the flow. The tool I employed in this work is a computer program that predicts broadband noise from fans. The program assumes that the complex shape of the curved blade can be represented as a single flat plate, allowing it to use fairly simple equations that can be solved in a reasonable amount of time. While the results from such representation provided reasonable estimates of the broadband noise levels, they did not usually represent the entire spectrum accurately. My investigation found that the discrepancy between data and theory can be improved if the leading edge and the trailing edge of the blade are treated separately. Using this approach, I reduced the maximum error in noise level from a high of 30% to less than 5% for the cases investigated. Detailed results of this investigation will be discussed at my presentation. The objective of this study is to investigate the influence of geometric parameters

Unton, Timothy J.↗

NASA NTRS1973

Concepts for a theoretical and experimental study of lifting rotor random loads and vibrations (identification of lifting rotor system parameters from transient response data), Phase 7-B

System identification methods have been applied to rotorcraft to estimate stability derivatives from transient flight control response data. While these applications assumed a linear constant coefficient representation of the rotorcraft, the computer experiments used transient responses in flap-bending and torsion of a rotor blade at high advance ratio which is a rapidly time varying periodic system. It was found that a simple system identification method applying a linear sequential estimator also called least square estimator or equation of motion estimator, is suitable for this periodic system and can be used directly if only the acceleration data are noise polluted. In the case of noise being present also in the state variable data the direct application of the estimator gave poor results.

Hohenemser, K. H.↗

NASA NTRS1980

Fan noise caused by the ingestion of anisotropic turbulence - A model based on axisymmetric turbulence theory

An analytical model of fan noise caused by inflow turbulence, a generalization of earlier work by Mani, is presented. Axisymmetric turbulence theory is used to develop a statistical representation of the inflow turbulence valid for a wide range of turbulence properties. Both the dipole source due to rotor blade unsteady forces and the quadrupole source resulting from the interaction of the turbulence with the rotor potential field are considered. The effects of variations in turbulence properties and fan operating conditions are evaluated. For turbulence axial integral length scales much larger than the blade spacing, the spectrum is shown to consist of sharp peaks at the blade passing frequency and its harmonics, with negligible broadband content. The analysis can then be simplified considerably and the total sound power contained within each spectrum peak becomes independent of axial length scale, while the width of the peak is inversely proportional to this parameter. Large axial length scales are characteristic of static fan test facilities, where the transverse contraction of the inlet flow produces highly anisotropic turbulence. In this situation, the rotor/turbulence interaction noise is mainly caused by the transverse component of turbulent velocity.

Kerschen, E. J.↗

NASA NTRS1982

A mathematical model of a single main rotor helicopter for piloted simulation

A mathematical model, suitable for piloted simulation of the flying qualities of helicopters, is a nonlinear, total force and moment model of a single main rotor helicopter. The model has ten degrees of freedom: six rigid body, three rotor flapping, and the rotor rotational degrees of freedom. The rotor model assumes rigid blades with rotor forces and moments radially integrated and summed about the azimuth. The fuselage aerodynamic model uses a detailed representation over a nominal angle of attack and sideslip range of + or - 15 deg., as well as a simplified curve fit at large angles of attack or sideslip. Stabilizing surface aerodynamics are modeled with a lift curve slope between stall limits and a general curve fit for large angles of attack. A generalized stability and control augmentation system is described. Additional computer subroutines provide options for a simplified engine/governor model, atmospheric turbulence, and a linearized six degree of freedom dynamic model for stability and control analysis.

Talbot, P. D.↗