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At least 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Aeroelastic Analysis of the NASA/ARMY/MIT Active Twist Rotor

Aeroelastic modeling procedures used in the design of a piezoelectric controllable twist helicopter rotor wind tunnel model are described. Two aeroelastic analysis methods developed for active twist rotor studies, and used in the design of the model blade, are described in this paper. The first procedure uses a simple flap-torsion dynamic representation of the active twist blade, and is intended for rapid and efficient control law and design optimization studies. The second technique employs a commercially available comprehensive rotor analysis package, and is used for more detailed analytical studies. Analytical predictions of hovering flight twist actuation frequency responses are presented for both techniques. Forward flight fixed system nP vibration suppression capabilities of the model active twist rotor system are also presented. Frequency responses predicted using both analytical procedures agree qualitatively for all design cases considered, with best correlation for cases where uniform blade properties are assumed.

Wilkie, W. Keats↗
NASA NTRS1985

Coolant passage heat transfer with rotation

The objective is to develop a heat transfer and pressure drop data base, computational fluid dynamic techniques, and correlations for multi-pass rotating coolant passages with and without flow turbulators. The experimental effort is focused on the simulation of configurations and conditions expected in the blades of advanced aircraft high pressure turbines. With the use of this data base, the effects of Coriolis and buoyancy forces on the coolant side flow can be included in the design of turbine blades.

Hajek, T. J.↗
NASA NTRS1986

Coolant passage heat transfer with rotation

In current and advanced gas turbine engines, increased speeds, pressures and temperatures are used to reduce specific fuel consumption and increase thrust/weight ratios. Hence, the turbine airfoils are subjected to increased heat loads escalating the cooling requirements to satisfy life goals. The efficient use of cooling air requires that the details of local geometry and flow conditions be adequately modeled to predict local heat loads and the corresponding heat transfer coefficients. The objective of this program is to develop a heat transfer and pressure drop data base, computational fluid dynamic techniques and correlations for multi-pass rotating coolant passages with and without flow turbulators. The experimental effort is focused on the simulation of configurations and conditions expected in the blades of advanced aircraft high pressure turbines. With the use of this data base, the effects of Coriolis and buoyancy forces on the coolant side flow can be included in the design of turbine blades.

Hajek, T. J.↗
NASA NTRS1987

Coolant passage heat transfer with rotation

The objective is to develop a heat transfer and pressure drop data base, computational fluid dynamic techniques and heat transfer correlations for rotating multipass coolant passages, with and without flow tabulators. The experimental effort is focused on the simulation of configurations and conditions expected in the blades of advanced aircraft high pressure turbines. With the use of this data base, the effects of Coriolis and buoyancy forces on the coolant side flow can be included in the design of turbine blades.

Hajek, T. J.↗
NASA NTRS1993

Evaluation of MARC for the analysis of rotating composite blades

The suitability of the MARC code for the analysis of rotating composite blades was evaluated using a four-task process. A nonlinear displacement analysis and subsequent eigenvalue analysis were performed on a rotating spring mass system to ensure that displacement-dependent centrifugal forces were accounted for in the eigenvalue analysis. Normal modes analyses were conducted on isotropic plates with various degrees of twist to evaluate MARC's ability to handle blade twist. Normal modes analyses were conducted on flat composite plates to validate the newly developed coupled COBSTRAN-MARC methodology. Finally, normal modes analyses were conducted on four composite propfan blades that were designed, analyzed, and fabricated at NASA Lewis Research Center. Results were compared with experimental data. The research documented herein presents MARC as a viable tool for the analysis of rotating composite blades.

Bartos, Karen F.↗
NASA NTRS1986

On the application of a linearized unsteady potential-flow analysis to fan-tip cascades

A linearized potential flow analysis, which accounts for the effects of nonuniform steady flow phenomena on the unsteady response to prescribed blade motions, has been applied to five two-dimensional cascade configurations. These include a flat-plate cascade and three cascades which are representative of the tip sections of current fan designs. Here the blades are closely spaced, highly staggered, and operate at low mean incidence. The fifth configuration is a NASA Lewis cascade of symmetric biconvex airfoils for which experimental measurements are available. Numerical solutions are presented that clearly illustrate the effects and importance of blade geometry and mean blade loading on the linearized unsteady response at high subsonic inlet Mach number and high blade-vibrational frequency. In addition, a good qualitative agreement is shown between the analytical predictions and experimental measurements for the cascade of symmetric biconvex airfoils. Finally, recommendations on the research needed to extend the range of application of linearized unsteady aerodynamic analyses are provided.

Usab, W. J., Jr.↗
NASA NTRS1981

Testing and evaluation of a stall-flutter-suppression system for helicopter rotors using individual-blade-control

The development and testing of a feedback system designed to alleviate the violent blade first torsion mode oscillations associated with stall flutter are described. The system, based on previously developed M.I.T. Individual-Blade-Control hardware, employs blade-mounted accelerometers to sense torsional oscillations and feeds back rate informaton to increase the damping of the first torsion mode. A linear model of the blade and control system dynamics is developed and is used to give qualitative and quantitative guidance in the design process as well as to aid in analysis of experimental results. System performance in wind tunnel tests, both in hover and forward flight, is described, and evidence is given of the system's ability to provide substantial additional damping to stall-induced blade oscillations.

Quackenbush, T. R.↗
NASA NTRS1986

Probabilistic structural analysis methods for critical SSME propulsion components

The development of a three-dimensional inelastic analysis methodology for the Space Shuttle main engine (SSME) structural components is described. The methodology is composed of: (1) composite load spectra, (2) probabilistic structural analysis methods, (3) the probabilistic finite element theory, and (4) probabilistic structural analysis. The progress in the development of generic probabilistic models for various individual loads which consist of a steady state load, a periodic load, a random load, and a spike, is discussed. The capabilities of the Numerical Evaluation of Stochastic Structures Under Stress finite element code designed for probabilistic structural analysis of the SSME are examined. Variation principles for formulation probabilistic finite elements and a structural analysis for evaluating the geometric and material properties tolerances on the structural response of turbopump blades are being designed.

Chamis, C. C.↗
NASA NTRS2011

Application of Probabilistic Methods to Assess Risk Due to Resonance in the Design of J-2X Rocket Engine Turbine Blades

The LOX-Hydrogen J-2X Rocket Engine, which is proposed for use as an upper-stage engine for numerous earth-to-orbit and heavy lift launch vehicle architectures, is presently in the design phase and will move shortly to the initial development test phase. Analysis of the design has revealed numerous potential resonance issues with hardware in the turbomachinery turbine-side flow-path. The analysis of the fuel pump turbine blades requires particular care because resonant failure of the blades, which are rotating in excess of 30,000 revolutions/minutes (RPM), could be catastrophic for the engine and the entire launch vehicle. This paper describes a series of probabilistic analyses performed to assess the risk of failure of the turbine blades due to resonant vibration during past and present test series. Some significant results are that the probability of failure during a single complete engine hot-fire test is low (~1%) because of the small likelihood of resonance, but that the probability increases to around 30% for a more focused turbomachinery-only test because all speeds will be ramped through and there is a greater likelihood of dwelling at more speeds. These risk calculations have been invaluable for use by program management in deciding if risk-reduction methods such as dampers are necessary immediately or if the test can be performed before the risk-reduction hardware is ready.

Brown, Andrew M.↗
NASA NTRSDate not supplied

Optimizing Proprotor Blades Using Coupled Aeroacoustic and Aerodynamic Sensitivities

A quieter and aerodynamically more efficient proprotor design requires high-fidelity and well-integrated optimization and analysis tools. To fulfill that requirement, the present paper delivers a methodology based on multidisciplinary, adjoint-based, discrete optimization. SU2-based code development involves the implementation of aeroacoustic analysis, adjoint computations, and integrations into a multidisciplinary rotorcraft optimization suite. Submodules utilized in the optimization are verified with wind tunnel data to demonstrate the accuracy of aerodynamic and aeroacoustic analyses. The developed code is used for NASA's helically twisted proprotor to maximize the aeroacoustic performance of the proprotor while holding thrust constant. The optimization process considers multiple flight conditions (hence, multipoint), which are forward flight and hovering. As an outcome of the analyses, the optimized blade design propagates lower noise as perceived by multiple observers in both flight conditions

R Omur Icke↗
NASA NTRS2024

Forced Response Aeromechanics Analysis in MATLAB®-Based Environment Code With Application to Distortion-Tolerant Fan R24 Blade Geometry

Researchers at the NASA Glenn Research Center have developed a research-level code for forced response analysis for turbomachinery that provides a streamlined framework for aeromechanics analysis, as well as for generating Goodman diagrams. The Forced Response Aeromechanics Analysis in a MATLAB-Based Environment (FRAAME, Version 1) code is designed to accept blade surface unsteady pressure time histories generated by three-dimensional (3D) Unsteady Reynolds-Averaged Navier-Stokes (URANS) code TURBO and modal displacements, modal stresses, and static stresses generated via expanded ANSYS® (Ansys, Inc.) cyclic symmetry analysis. The code’s looping structure allows for largescale analyses, including many blade surface pressure files, modal displacements, and modal stress files for full annulus forced response analysis, including a modal summation method for multiple engine orders and modes. This code is applied to the R24 fan geometry for the Boundary Layer Ingesting Inlet/Distortion-Tolerant Fan (BLI2DTF) project, a propulsion system being developed to increase the fuel efficiency of future aircraft, to serve as validation of experimental data as well as external simulation results. This tool’s functional, modular form is intended to allow users to make modifications easily. Those modifications can include requests for information for any step in the analysis process, as well as adding various functions to compute additional information of interest. It is also written, given sufficient input information, as a general blade geometry forced response solver that is not necessarily specific to the R24 geometry but is used here for verification of code functionality and accuracy. Although forced response analysis tools are also available in commercially available code packages, the post-processing FRAAME code offers reliable, fast, and scalable forced response analysis and Goodman diagram generation for large cases utilizing, specifically, TURBO and ANSYS® results information in an effort to streamline forced response and high cycle fatigue analyses.

Aeromechanics↗
NASA NTRS1976

Aerodynamic performance of a 1.20-pressure ratio fan stage designed for low noise

The aerodynamic design and the overall blade element performance of a 51 centimeter diameter fan stage is presented. The stage was designed to minimize the noise generated by rotor stator interactions. The design pressure ratio was 1.20 at a flow of 30.6 kilograms per second and a rotor blade tip speed of 228.6 meters per second. At design speed the rotor peak efficiency was 0.935. The peak efficiency of the stage, however, was 0.824. The radial distribution of rotor performance parameters at peak efficiency and design speed indicated excellent agreement with design values.

Lewis, G. W., Jr.↗
NASA NTRS1944

Tests of Thermal-electric De-icing Equipment for Propellers

Flights were made in natural icing conditions at the NACA Ice Research Project, Minneapolis, Minn. to test several designs of thermal-electric propeller de-icing blade shoes and a hub-generator design. It was found that a minimum average unit power of 2.5 watts per square inch of blade-shoe area would protect the propeller blades at the test conditions. The most satisfactory blade shoe of the three designs tested extended to the 20-percent-chord point and to 90 percent of the blade radius. A concentration of heat in the leading-edge region of this shoe was found to reduce the power input necessary for satisfactory de-icing. A satisfactory thermal design of blade shoe and a hub generator of sufficient capacity were developed.

PROPELLERS - HAMILTON STANDARD↗
NASA NTRS1993

Forced response of mistuned bladed disk assemblies

A complete analytic model of mistuned bladed disk assemblies, designed to simulate the dynamical behavior of these systems, is analyzed. The model incorporates a generalized method for describing the mistuning of the assembly through the introduction of specific mistuning modes. The model is used to develop a computational bladed disk assembly model for a series of parametric studies. Results are presented demonstrating that the response amplitudes of bladed disk assemblies depend both on the excitation mode and on the mistune mode.

Watson, Brian C.↗
NASA NTRS1980

Superhybrid composite blade impact studies

The feasibility of superhybrid composite blades for meeting the mechanical design and impact resistance requirements of large fan blades for aircraft turbine engine applications was investigated. Two design concepts were evaluated: leading edge spar (TiCom) and center spar (TiCore), both with superhybrid composite shells. The investigation was both analytical and experimental. The results obtained show promise that superhybrid composites can be used to make light weight, high quality, large fan blades with good structural integrity. The blades tested successfully demonstrated their ability to meet steady state operating conditions, overspeed, and small bird impact requirements.

Chamis, C. C.↗
NASA NTRS1959

Experimental Investigation of a 0.35 Hub-Tip Radius Ratio Transonic Axial Flow Rotor Designed for 40 Pounds per Second per Square Foot with a Design Tip Diffusion Factor of 0.20

In order to determine the effect of a low design diffusion factor on the performance of a transonic axial-flow compressor rotor, a high-specific-flow rotor with a 0.35 hub-tip radius ratio was designed, fabricated and tested. This rotor used a design tip diffusion factor of 0.20 with a design corrected specific weight flow of 40 pounds per second per square foot of frontal area, a total-pressure ratio of 1.27, and an adiabatic efficiency of 0.96. The design, rotor performance, and blade element performance are presented with a discussion on rotor shock losses and a comparison with a similarly designed rotor with a tip diffusion factor of 0.35. At the design corrected tip speed of 1100 feet per second, a peak rotor adiabatic efficiency of 0.88 was attained at a corrected specific weight flow of 39 pounds per second per square foot of frontal area with a mass-averaged total-pressure ratio of 1.27. The blade element tip diffusion factor was 0.281, which is 0.08 higher than the design value of 0.20. Peak efficiencies of 0.95, 0.91, 0.89, and 0.85 were obtained at 70, 80, 90, and 110 percent of design speed, respectively. Comparison of the performance of the rotor reported herein and a similarly designed rotor with increased blade loading indicates that higher blade loading results in a more desirable rotor because of a higher pressure ratio and equivalent efficiency. Computed values of shock losses at the rotor tip section indicate that the losses at peak efficiency are primarily a function of shock losses since the profile losses are only a small percentage of the total loss.

Yasaki, Paul T.↗
NASA NTRS1985

Oscillating airfoils and their wake

The unsteady phenomena in the wake of an oscillating wing or rotor blade are examined theoretically using the Prandtl approximation of the vortex-transport equation. A mathematical model is developed and applied to such problems as the effect of winglets on the performance of fixed wings and the possibly of employing similar designs in rotor blades. Model predictions for several profiles are compared with published and experimental measurements, and good agreement is found. Graphs and diagrams are provided.

Send, W.↗
NASA NTRS1982

Response studies of rotors and rotor blades with application to aeroelastic tailoring

Various tools for the aeroelastic stability and response analysis of rotor blades in hover and forward flight were developed and incorporated in a comprehensive package capable of performing aeroelastic tailoring of rotor blades in forward flight. The results indicate that substantial vibration reductions, of order 15-40%, in the vibratory hub shears can be achieved by relatively small modifications of the initial design. Furthermore the optimized blade can be up to 20% lighter than the original design. Accomplishments are reported for the following tasks: (1) finite element modeling of rotary-wing aeroelastic problems in hover and forward flight; (2) development of numerical methods for calculating the aeroelastic response and stability of rotor blades in forward fight; (3) formulation of the helicopter air resonance problem in hover with active controls; and (4) optimum design of rotor blades for vibration reduction in forward flight.

Friedmann, P. P.↗