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At least 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Dynamic analysis of a pre-and-post ice impacted blade

The dynamic characteristics of an engine blade are evaluated under pre-and-post ice impact conditions using the NASA in-house computer code BLASIM. The ice impacts the leading edge of the blade causing severe local damage. The local structural response of the blade due to the ice impact is predicted via a transient response analysis by modeling only a local patch around the impact region. After ice impact, the global geometry of the blade is updated using deformations of the local patch and a free vibration analysis is performed. The effects of ice impact location, size and ice velocity on the blade mode shapes and natural frequencies are investigated. The results indicate that basic nature of the mode shapes remains unchanged after impact and that the maximum variation in natural frequencies occurs for the twisting mode of the blade.

Abumeri, G. H.↗
NASA NTRS1990

Two- and three-dimensional blade vortex interactions

A three-dimensional time dependent Navier-Stokes analysis was applied to the rotor blade vortex interaction (BVI) problem. The numerical procedure is an iterative implicit procedure using three point central differences to represent spatial derivatives. A series of calculations were made to determine the time steps, pseudo-time steps, iterations, artificial dissipation level, etc. required to maintain a nondissipative vortex. Results show the chosen method to have excellent non-dissipative properties provided the correct parameters are chosen. This study was used to set parameters for both two- and three-dimensional blade vortex interaction studies. The two-dimensional study considered the interaction between a vortex and a NACA0012 airfoil. The results showed the detailed physics during the interaction including the pressure pulse propagating from the blade. The simulated flow physics was qualitatively similar to that experimentally observed. The 2-D BVI phenomena is the result of the buildup and violent collapse of the shock waves and local supersonic pockets on the blade surfaces. The resulting pressure pulse build-up appears to be centered at the blade leading edge. The three-dimensional interaction study considered the case of a vortex at 20 deg incidence to the blade leading edge. Although the qualitative results were similar to that of the two-dimensional interaction, details clearly showed the three-dimensional nature of the interaction process.

Davoudzadeh, F.↗
NASA NTRS1993

Blade Assessment for Ice Impact (BLASIM). User's manual, version 1.0

The Blade Assessment Ice Impact (BLASIM) computer code can analyze solid, hollow, composite, and super hybrid blades. The solid blade is made up of a single material where hollow, composite, and super hybrid blades are constructed with prescribed composite layup. The properties of a composite blade can be specified by inputting one of two options: (1) individual ply properties, or (2) fiber/matrix combinations. When the second option is selected, BLASIM utilizes ICAN (Integrated Composite ANalyzer) to generate the temperature/moisture dependent ply properties of the composite blade. Two types of geometry input can be given: airfoil coordinates or NASTRAN type finite element model. These features increase the flexibility of the program. The user's manual provides sample cases to facilitate efficient use of the code while gaining familiarity.

Reddy, E. S.↗
NASA NTRS1992

Flow visualizations of perpendicular blade vortex interactions

Helium bubble flow visualizations have been performed to study perpendicular interaction of a turbulent trailing vortex and a rectangular wing in the Virginia Tech Stability Tunnel. Many combinations of vortex strength, vortex-blade separation (Z(sub s)) and blade angle of attack were studied. Photographs of representative cases are presented. A range of phenomena were observed. For Z(sub s) greater than a few percent chord the vortex is deflected as it passes the blade under the influence of the local streamline curvature and its image in the blade. Initially the interaction appears to have no influence on the core. Downstream, however, the vortex core begins to diffuse and grow, presumably as a consequence of its interaction with the blade wake. The magnitude of these effects increases with reduction in Z(sub s). For Z(sub s) near zero the form of the interaction changes and becomes dependent on the vortex strength. For lower strengths the vortex appears to split into two filaments on the leading edge of the blade, one passing on the pressure and one passing on the suction side. At higher strengths the vortex bursts in the vicinity of the leading edge. In either case the core of its remnants then rapidly diffuse with distance downstream. Increase in Reynolds number did not qualitatively affect the flow apart from decreasing the amplitude of the small low-frequency wandering motions of the vortex. Changes in wing tip geometry and boundary layer trip had very little effect.

Rife, Michael C.↗
NASA NTRS1993

Fabrication of composite propfan blades for a cruise missile wind tunnel model

This report outlines the procedures that were employed in fabricating prototype graphite-epoxy composite prop fan blades. These blades were used in wind tunnel tests that investigated prop fan propulsion system interactions with a missile airframe in order to study the feasibility of an advanced-technology-propfan-propelled missile. Major phases of the blade fabrication presented include machining of the master blade, mold fabrication, ply cutting and assembly, blade curing, and quality assurance. Specifically, four separate designs were fabricated, 18 blades of each geometry, using the same fabrication technique for each design.

Fite, E. Brian↗
NASA NTRS1992

Dynamic analysis of a pre-and-post ice impacted blade

The dynamic characteristics of an engine blade are evaluated under pre-and-post ice impact conditions using the NASA in-house computer code BLASIM. The ice impacts the leading edge of the blade causing severe local damage. The local structural response of the blade due to the ice impact is predicted via a transient response analysis by modeling only a local patch around the impact region. After ice impact, the global geometry of the blade is updated using deformations of the local patch and a free vibration analysis is performed. The effects of ice impact location, ice size and ice velocity on the blade mode shapes and natural frequencies are investigated. The results indicate that basic nature of the mode shapes remains unchanged after impact and that the maximum variation in natural frequencies occurs for the twisting mode of the blade.

Abumeri, G. H.↗
NASA NTRS1992

Advanced Technology Blade testing on the XV-15 Tilt Rotor Research Aircraft

The XV-15 Tilt Rotor Research Aircraft has just completed the first series of flight tests with the Advanced Technology Blade (ATB) rotor system. The ATB are designed specifically for flight research and provide the ability to alter blade sweep and tip shape. A number of problems were encountered from first installation through envelope expansion to airplane mode flight that required innovative solutions to establish a suitable flight envelope. Prior to operation, the blade retention hardware had to be requalified to a higher rated centrifugal load, because the blade weight was higher than expected. Early flights in the helicopter mode revealed unacceptably high vibratory control system loads which required a temporary modification of the rotor controls to achieve higher speed flight and conversion to airplane mode. The airspeed in airplane mode was limited, however, because of large static control loads. Furthermore, analyses based on refined ATB blade mass and inertia properties indicated a previously unknown high-speed blade mode instability, also requiring airplane-mode maximum airspeed to be restricted. Most recently, a structural failure of an ATB cuff (root fairing) assembly retention structure required a redesign of the assembly. All problems have been addressed and satisfactory solutions have been found to allow continued productive flight research of the emerging tilt rotor concept.

Wellman, Brent↗
NASA NTRS1993

Optimization of blade arrangement in a randomly mistuned cascade using simulated annealing

This paper presents preliminary results of an investigation on mistuning of bladed-disk assemblies aimed at capturing the benefits of mistuning on stability, while at the same time, minimizing the adverse effects on response by solving the following problem: given a set of N turbine blades, each being a small random perturbation of the same nominal blade, determine the best arrangement of the N blades in a mistuned cascade with regard to aeroelastic response. In the studies reported here, mistuning of the blades is restricted to small differences in torsional stiffness. The large combinatorial optimization problem of seeking the best arrangement by blade exchanges is solved using a simulated annealing algorithm.

Thompson, Edward A.↗
NASA NTRS1951

Two-dimensional Cascade Investigation of the Maximum Exit Tangential Velocity Component and Other Flow Conditions at the Exit of Several Turbine Blade Designs at Supercritical Pressure Ratios

The nature of the flow at the exit of a row of turbine blades for the range of conditions represented by four different blade configurations was evaluated by the conservation-of-momentum principle using static-pressure surveys and by analysis of Schlieren photographs of the flow. It was found that for blades of the type investigated, the maximum exit tangential-velocity component is a function of the blade geometry only and can be accurately predicted by the method of characteristics. A maximum value of exit velocity coefficient is obtained at a pressure ratio immediately below that required for maximum blade loading followed by a sharp drop after maximum blade loading occurs.

TURBINES - AXIAL FLOW↗
NASA NTRS1952

Experimental investigation of the heat-transfer characteristics of an air-cooled sintered porous turbine blade

The results of a heat-transfer investigation on a transpiration-cooled turbine blade are presented. The investigation was made on a static cascade.506::Declassified December 11, 1953. Blade-peripheral wall-temperature distributions were measured and compared with the temperature distribution on a convection-cooled blade. A correlation of the blade-wall temperatures obtained at various gas temperatures and gas weight flows was made on the basis of a temperature-difference ratio. Blade-wall temperatures calculated using an approximate theory for turbulent flow are compared with the measured blade-wall tempeartures.

Schafer, Louis J , Jr↗
NASA NTRS1956

Analysis of turbomachine viscous losses affected by changes in blade geometry

The effect of changes in blade geometry on the viscous losses in turbomachines is analyzed. The viscous losses can be expressed as a function of three variables: the blade height-to spacing ratio, the solidity, and a blade-height Reynolds number. A method of optimizing the number of blades and the solidity is given. The effect of trailing-edge blockage on blade loss and its effect on the choice of blade number for a particular design are also discussed.

Miser, James W↗
NASA NTRS1950

Investigation of the NACA 4-(3)(8)-045 Two-blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Compressibility and Solidity on Performance

As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.

Stack, John↗
NASA NTRS1958

Effect of Chord Size on Weight and Cooling Characteristics of Air-Cooled Turbine Blades

An analysis has been made to determine the effect of chord size on the weight and cooling characteristics of shell-supported, air-cooled gas-turbine blades. In uncooled turbines with solid blades, the general practice has been to design turbines with high aspect ratio (small blade chord) to achieve substantial turbine weight reduction. With air-cooled blades, this study shows that turbine blade weight is affected to a much smaller degree by the size of the blade chord.

Esgar, Jack B↗
NASA NTRS1993

Blading models for TURBAN and CSPAN turbomachine design codes

Blading models added to TURBAN, an axial turbine meanline design code, and CSPAN, an axial compressor spanline design code, are presented. Calculations of blade chord length, blade axial length, and number of blades for each blade row are now provided by each code. In order to accomplish this, computational models for axial solidity and stagger angle were added to TURBAN. For a compressor, where solidity and chord were already available from input, the modeling additions to CSPAN were for stagger angle and for the blade angles to include incidence and deviation. All added computations utilized models already available in the literature.

Glassman, Arthur J.↗
NASA NTRS1993

Simulation of unsteady flows through stator and rotor blades of a gas turbine using the Chimera method

A two-dimensional model to solve compressible Navier-Stokes equations for the flow through stator and rotor blades of a turbine is developed. The flow domains for the stator and rotor blades are coupled by the Chimera method that makes grid generation easy and enhances accuracy because the area of the grid that have high turning of grid lines or high skewness can be eliminated from the computational domain after the grids are generated. The results of flow computations show various important features of unsteady flows including the acoustic waves interacting with boundary layers, Karman vortex shedding from the trailing edge of the stator blades, pulsating incoming flow to a rotor blade from passing stator blades, and flow separation from both suction and pressure sides of the rotor blades.

Nakamura, S.↗
NASA NTRS1994

Forward sweep, low noise rotor blade

A forward-swept, low-noise rotor blade includes an inboard section, an aft-swept section, and a forward-swept outboard section. The rotor blade reduces the noise of rotorcraft, including both standard helicopters and advanced systems such as tiltrotors. The primary noise reduction feature is the forward sweep of the planform over a large portion of the outer blade radius. The rotor blade also includes an aft-swept section. The purpose of the aft-swept region is to provide a partial balance to pitching moments produced by the outboard forward-swept portion of the blade. The noise source showing maximum noise reduction is blade-vortex interaction (BVI) noise. Also reduced are thickness, noise, high speed impulsive noise, cabin vibration, and loading noise.

Brooks, Thomas F.↗
NASA NTRS1992

Aeroelastic modeling of composite rotor blades with straight and swept tips

This paper presents an analytical study of the aeroelastic behavior of composite rotor blades with straight and swept tips. The blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. The nonlinear equations of motion for the FEM are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. It is shown that composite ply orientation has a substantial effect on blade stability. At low thrust conditions, certain ply orientations can cause instability in the lag mode. The flap-torsion coupling associated with tip sweep can also induce aeroelastic instability in the blade. This instability can be removed by appropriate ply orientation in the composite construction. These results illustrate the inherent potential for aeroelastic tailoring present in composite rotor blades with swept tips, which still remains to be exploited in the design process.

Yuan, Kuo-An↗
NASA NTRS1995

Testing And Analysis Of Rubbing Of Turbine-Blade Tips

Unique combination of established techniques of experimentation and theoretical analysis devised to study stresses induced in turbine blades by intermittent or periodic rubbing of tips of blades in turbine housings. Rubbing is inevitable consequence of modern high-speed, high-efficiency turbine designs, which call for blade-tip clearances as small as possible. Necessary to quantify stresses caused by rubbing in order to predict fatigue lives of blades. Testing apparatus simulates blade-tip rubbing and measures dynamic responses of blades.

Davis, Gary A.↗