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At least 559 records · Page 31

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NASA NTRS1999

Successful Solutions to SSME/AT Development Turbine Blade Distress

As part of the High-Pressure Fuel Turbopump/Alternate Turbopump (HPFTP/AT) turbine blade development program, unique turbine blade design features were implemented to address 2nd stage turbine blade high cycle fatigue distress and improve turbine robustness. Features included the addition of platform featherseal dampers, asymmetric blade tip seal segments, gold plating of the blade attachments, and airfoil tip trailing edge modifications. Development testing shows these features have eliminated turbine blade high cycle fatigue distress and consequently these features are currently planned for incorporation to the flight configuration. Certification testing will begin in 1999. This presentation summarizes these features.

Montgomery, Stuart K.↗

NASA NTRS1999

Aero-Thermo-Structural Design Optimization of Internally Cooled Turbine Blades

A set of robust and computationally affordable inverse shape design and automatic constrained optimization tools have been developed for the improved performance of internally cooled gas turbine blades. The design methods are applicable to the aerodynamics, heat transfer, and thermoelasticity aspects of the turbine blade. Maximum use of the existing proven disciplinary analysis codes is possible with this design approach. Preliminary computational results demonstrate possibilities to design blades with minimized total pressure loss and maximized aerodynamic loading. At the same time, these blades are capable of sustaining significantly higher inlet hot gas temperatures while requiring remarkably lower coolant mass flow rates. These results suggest that it is possible to design internally cooled turbine blades that will cost less to manufacture, will have longer life span, and will perform as good, if not better than, film cooled turbine blades.

Dulikravich, G. S.↗

NASA NTRS1997

Adiabatic Effectiveness and Heat Transfer Coefficient on a Film-Cooled Rotating Blade

three-dimensional Navier-Stokes code has been used to compute the adiabatic effectiveness and heat transfer coefficient on a rotating film-cooled turbine blade. The blade chosen is the United Technologies Research Center(UTRC) rotor with five film-cooling rows containing 83 holes, including three rows on the shower head with 49 holes, covering about 86% of the blade span. The mainstream is akin to that under real engine conditions with stagnation temperature 1900 K and stagnation pressure 3 MPa. The blade speed is taken to be 5200 rpm. The adiabatic effectiveness is higher for a rotating blade as compared to that for a stationary blade. Also, the direction of coolant injection from the shower-head holes considerably affects the effectiveness and heat transfer coefficient values on both the pressure and suction surfaces. In all cases the heat transfer coefficient and adiabatic effectiveness are highly three-dimensional in the vicinity of holes but tend to become two-dimensional far downstream.

Garg, Vijay K.↗

NASA NTRS2000

A Novel Method for Reducing Rotor Blade-Vortex Interaction

One of the major hindrances to expansion of the rotorcraft market is the high-amplitude noise they produce, especially during low-speed descent, where blade-vortex interactions frequently occur. In an attempt to reduce the noise levels caused by blade-vortex interactions, the flip-tip rotor blade concept was devised. The flip-tip rotor increases the miss distance between the shed vortices and the rotor blades, reducing BVI noise. The distance is increased by rotating an outboard portion of the rotor tip either up or down depending on the flight condition. The proposed plan for the grant consisted of a computational simulation of the rotor aerodynamics and its wake geometry to determine the effectiveness of the concept, coupled with a series of wind tunnel experiments exploring the value of the device and validating the computer model. The computational model did in fact show that the miss distance could be increased, giving a measure of the effectiveness of the flip-tip rotor. However, the wind experiments were not able to be conducted. Increased outside demand for the 7'x lO' wind tunnel at NASA Ames and low priority at Ames for this project forced numerous postponements of the tests, eventually pushing the tests beyond the life of the grant. A design for the rotor blades to be tested in the wind tunnel was completed and an analysis of the strength of the model blades based on predicted loads, including dynamic forces, was done.

Glinka, A. T.↗

NASA NTRS1994

Reduction of Helicopter BVI Noise, Vibration, and Power Consumption Through Individual Blade Control

A wind tunnel test was conducted with a full-scale BO 105 helicopter rotor to evaluate the potential of open-loop individual blade control (IBC) to improve rotor performance, to reduce blade vortex interaction (BVI) noise, and to alleviate helicopter vibrations. The wind tunnel test was an international collaborative effort between NASA/U.S. Army AFDD, ZF Luftfahrttechnik, Eurocopter Deutschland, and the German Aerospace Laboratory (DLR) and was conducted under the auspices of the U.S./German MOU on Rotorcraft Aeromechanics. In this test the normal blade pitch links of the rotor were replaced by servo-actuators so that the pitch of each blade could be controlled independently of the other blades. The specially designed servoactuators and IBC control system were designed and manufactured by ZF Luftfahrttechnik, GmbH. The wind tunnel test was conducted in the 40- by 80-Foot Wind Tunnel at the NASA Ames Research Center. An extensive amount of measurement information was acquired for each IBC data point. These data include rotor performance, static and dynamic hub forces and moments, rotor loads, control loads, inboard and outboard blade pitch motion, and BVI noise data. The data indicated very significant (80 percent) simultaneous reductions in both BVI noise and hub vibrations could be obtained using multi-harmonic input at the critical descent (terminal approach) condition. The data also showed that performance improvements of up to 7 percent could be obtained using 2P input at high-speed forward flight conditions.

Jacklin, Stephen A.↗

NASA NTRS1999

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub comer stall. Both the measurements and the numerical solutions for the stator indicate that a strong twister-like vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub comer stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Hah, C.↗

NASA NTRS1996

Active Control of Helicopter Blade Stall

In an effort to expand the helicopter flight envelope, this analytical study proposes the concept of using high frequency blade pitch actuation to alleviate blade stall at high speed and/or high thrust flight conditions. The availability of high-frequency blade-mounted actuators has made this concept realizable. This study is carried out using the University of Maryland Advanced Rotorcraft Code (or UMARC), which employs state-of-the-art structural and aerodynamic modelings. The salient features of this code include the application of the finite element methods in both space and time domains, and the incorporation of an advanced unsteady aerodynamic and nonuniform inflow models. Preliminary results indicate that two-per-rev blade pitch control can reduce retreating blade stall for a rotor operating at high speed and high thrust flight condition. Current efforts concentrate on developing an automatic stall suppression system which employs a combination of higher harmonic blade pitch schedule.

Nguyen, Khanh↗

NASA NTRS2003

Probabilistic Evaluation of Blade Impact Damage

The response to high velocity impact of a composite blade is probabilistically evaluated. The evaluation is focused on quantifying probabilistically the effects of uncertainties (scatter) in the variables that describe the impact, the blade make-up (geometry and material), the blade response (displacements, strains, stresses, frequencies), the blade residual strength after impact, and the blade damage tolerance. The results of probabilistic evaluations results are in terms of probability cumulative distribution functions and probabilistic sensitivities. Results show that the blade has relatively low damage tolerance at 0.999 probability of structural failure and substantial at 0.01 probability.

Chamis, C. C.↗

NASA NTRS2000

Hover Testing of the NASA/Army/MIT Active Twist Rotor Prototype Blade

Helicopter rotor individual blade control promises to provide a mechanism for increased rotor performance and reduced rotorcraft vibrations and noise. Active material methods, such as piezoelectrically actuated trailing-edge flaps and strain-induced rotor blade twisting, provide a means of accomplishing individual blade control without the need for hydraulic power in the rotating system. Recent studies have indicated that controlled strain induced blade twisting can be attained using piezoelectric active fiber composite technology. In order to validate these findings experimentally, a cooperative effort between NASA Langley Research Center, the Army Research Laboratory, and the MIT Active Materials and Structures Laboratory has been developed. As a result of this collaboration an aeroelastically-scaled active-twist model rotor blade has been designed and fabricated for testing in the heavy gas environment of the Langley Transonic Dynamics Tunnel (TDT). The results of hover tests of the active-twist prototype blade are presented in this paper. Comparisons with applicable analytical predictions of active-twist frequency response in hovering flight are also presented.

Matthew L Wilbur↗

NASA NTRS1998

Measurement of Rotorcraft Blade Deformation using Projection Moire Interferometry

Projection Moire Interferometry (PMI) has been used to obtain near instantaneous, quantitative blade deformation measurements of a generic rotorcraft model at several test conditions. These laser-based measurements provide quantitative, whole field, dynamic blade deformation profiles conditionally sampled as a function of rotor azimuth. The instantaneous nature of the measurements permits computation of the mean and unsteady blade deformation, blade bending, and twist. The PMI method is presented, and the image processing steps required to obtain quantitative deformation profiles from PMI interferograms are described. Experimental results are provided which show blade bending, twist, and unsteady motion. This initial proof-of-concept test has demonstrated the capability of PMI to acquire accurate, full field rotorcraft blade deformation data.

Fleming, Gary A.↗

NASA NTRS2002

Self-Tuning Impact Dampers Designed for Turbomachinery Blade Vibration Suppression

Turbomachinery blades are subject to aerodynamic forces that can lead to high-cycle-fatigue (HCF) failures. These failures will only increase as engineers begin to design blades without shrouds or as integrally bladed disks (blisks). These new designs will decrease blade damping significantly because the mechanical damping from shroud and blade joints will be eliminated. Also, it is difficult to design dampers for the engine environment with its extremely high centrifugal loads and high temperatures. The self-tuning impact damper has been designed to provide the additional damping required to avoid HCF while withstanding the harsh engine environment. In addition, the damper is placed within the engine blade itself rather than external to it.

Duffy, Kirsten P.↗

NASA NTRS2005

System Developed for Real-Time Blade-Flutter Monitoring in the Wind Tunnel

A real-time system has been developed to monitor flutter vibrations in turbomachinery. The system is designed for continuous processing of blade tip timing data at a rate of 10 MB/sec. A USB 2.0 interface provides uninterrupted real-time processing of the data, and the blade-tip arrival times are measured with a 50-MHz oscillator and a 24-bit pipelined architecture counter. The input stage includes a glitch catcher, which reduces the probability of detecting a ghost blade to negligible levels. A graphical user interface provides online interrogation of any blade tip from any light probe sensor. Alternatively, data from all blades and all sensors can be superimposed into a single composite scatter plot displaying the vibration amplitude of each blade.

Kurkov, Anatole P.↗

NASA NTRS1993

Counterrotating aircraft propulsor blades

A propulsor blade for an aircraft engine includes an airfoil section formed in the shape of a scimitar. A metallic blade spar is interposed between opposed surfaces of the blade and is bonded to the surfaces to establish structural integrity of the blade. The metallic blade spar includes a root end allowing attachment of the blade to the engine.

Nelson, Joey L.↗

NASA NTRS1988

Counterrotating aircraft propulsor blades

Nelson, Joey L.↗

NASA NTRS1957

Endurance Evaluation of Sintered, Porous, Strut-Supported Turbine Blades made by Federal-Mogul-Bower-Bearings, Incorporated, under Bureau of Aeronautics Contract NOas 55-124-C

Four strut-supported, transpiration-cooled turbine blades were investigated experimentally in a turbojet engine. The blade shells were fabricated by the mold-sintering method with spherical stainless-steel powder. Two blades were investigated in order to evolve suitable capping methods for the blade tip. Two other blades were used to evaluate the durability of the porous-shell material. The blades were investigated at a turbine-tip speed of 1305 feet per second, an average turbine-inlet temperature of about 1670 F, and at a porous-shell temperature limited to a maximum of approximately 1040 F.

Hickel, Robert O.↗

NASA NTRS2012

Blade Displacement Measurement Technique Applied to a Full-Scale Rotor Test

Blade displacement measurements using multi-camera photogrammetry were acquired during the full-scale wind tunnel test of the UH-60A Airloads rotor, conducted in the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel. The objectives were to measure the blade displacement and deformation of the four rotor blades as they rotated through the entire rotor azimuth. These measurements are expected to provide a unique dataset to aid in the development and validation of rotorcraft prediction techniques. They are used to resolve the blade shape and position, including pitch, flap, lag and elastic deformation. Photogrammetric data encompass advance ratios from 0.15 to slowed rotor simulations of 1.0, thrust coefficient to rotor solidity ratios from 0.01 to 0.13, and rotor shaft angles from -10.0 to 8.0 degrees. An overview of the blade displacement measurement methodology and system development, descriptions of image processing, uncertainty considerations, preliminary results covering static and moderate advance ratio test conditions and future considerations are presented. Comparisons of experimental and computational results for a moderate advance ratio forward flight condition show good trend agreements, but also indicate significant mean discrepancies in lag and elastic twist. Blade displacement pitch measurements agree well with both the wind tunnel commanded and measured values.

Abrego, Anita I.↗

NASA NTRS2012

Determination of HART I Blade Structural Properties by Laboratory Testing

The structural properties of higher harmonic Aeroacoustic Rotor Test (HART I) blades were measured using the original set of blades tested in the German-dutch wind tunnel (DNW) in 1994. the measurements include bending and torsion stiffness, geometric offsets, and mass and inertia properties of the blade. the measured properties were compared to the estimated values obtained initially from the blade manufacturer. The previously estimated blade properties showed consistently higher stiffness, up to 30 percent for the flap bending in the blade inboard root section.

Jung, Sung N.↗

NASA NTRS2013

Combined Experimental and Numerical Simulations of Thermal Barrier Coated Turbine Blades Erosion

A combined experimental and computational study was conducted to investigate the erosion of thermal barrier coated (TBC) blade surfaces by alumina particles ingestion in a single stage turbine. In the experimental investigation, tests of particle surface interactions were performed in specially designed tunnels to determine the erosion rates and particle restitution characteristics under different impact conditions. The experimental results show that the erosion rates increase with increased impingement angle, impact velocity and temperature. In the computational simulations, an Euler-Lagrangian two stage approach is used in obtaining numerical solutions to the three-dimensional compressible Reynolds Averaged Navier-Stokes equations and the particles equations of motion in each blade passage reference frame. User defined functions (UDF) were developed to represent experimentally-based correlations for particle surface interaction models which were employed in the three-dimensional particle trajectory simulations to determine the particle rebound characteristics after each surface impact. The experimentally based erosion UDF model was used to predict the TBC erosion rates on the turbine blade surfaces based on the computed statistical data of the particles impact locations, velocities and angles relative to the blade surface. Computational results are presented for the predicted TBC blade erosion in a single stage commercial APU turbine, for a NASA designed automotive turbine, and for the NASA turbine scaled for modern rotorcraft operating conditions. The erosion patterns in the turbines are discussed for uniform particle ingestion and for particle ingestion concentrated in the inner and outer 5 percent of the stator blade span representing the flow cooling the combustor liner.

Hamed, Awate↗