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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.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

A computational design method for transonic turbomachinery cascades

This paper describes a systematical computational procedure to find configuration changes necessary to modify the resulting flow past turbomachinery cascades, channels and nozzles, to be shock-free at prescribed transonic operating conditions. The method is based on a finite area transonic analysis technique and the fictitious gas approach. This design scheme has two major areas of application. First, it can be used for design of supercritical cascades, with applications mainly in compressor blade design. Second, it provides subsonic inlet shapes including sonic surfaces with suitable initial data for the design of supersonic (accelerated) exits, like nozzles and turbine cascade shapes. This fast, accurate and economical method with a proven potential for applications to three-dimensional flows is illustrated by some design examples.

Sobieczky, H.↗
NASA NTRS1986

Low-speed wind-tunnel tests of single- and counter-rotation propellers

A low-speed (Mach 0 to 0.3) wind-tunnel investigation was conducted to determine the basic performance, force and moment characteristics, and flow-field velocities of single- and counter-rotation propellers. Compared with the eight-blade single-rotation propeller, a four- by four- (4 x 4) blade counter-rotation propeller with the same blade design produced substantially higher thrust coefficients for the same blade angles and advance ratios. The results further indicated that ingestion of the wake from a supporting pylon for a pusher configuration produced no significant change in the propeller thrust performance for either the single- or counter-rotation propellers. A two-component laser velocimeter (LV) system was used to make detailed measurements of the propeller flow fields. Results show increasing slipstream velocities with increasing blade angle and decreasing advance ratio. Flow-field measurements for the counter-rotation propeller show that the rear propeller turned the flow in the opposite direction from the front propeller and, therefore, could eliminate the swirl component of velocity, as would be expected.

Dunham, D. M.↗
NASA NTRS1991

An aeroelastic analysis with a generalized dynamic wake

An aeroelastic model with generalized dynamic wake is developed for application in the integration of aerodynamic, dynamic, and structural optimization of a rotor blade. The investigation is carried out with special attention to efficiency and accuracy of aeroelastic modeling. Each blade is assumed to be an elastic beam undergoing flap bending, lead-lag bending, elastic twist and axial deflections. The nonuniform blade is discretized into finite beam elements, each of which consists of twelve degrees of freedom. Such important blade design variables as pretwist, and chordwise offsets of the blade center of gravity and of the aerodynamic center from the elastic axis have been included in the analysis. Aerodynamic loads are computed from unsteady blade element theory where the rotor three-dimensional unsteady wake is modeled using a generalized dynamic wake theory. The noncirculatory loads based on unsteady thin airfoil theory are also included.

He, Cheng J.↗
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 NTRS1979

The MOD-1 steel blade

The design, development, fabrication, testing, and transport of two 100 foot metal blades for the MOD-1 WTS are summarized. Because the metal blade design was started late in the MOD-1 system development, many of the design requirements (allocations) were restrictive for the metal blade concept, particularly the maximum weight requirement. The design solutions required to achieve the weight goal resulted in a labor intensive (expensive) fabrication, particularly for a quantity of only two blades manufactured using minimal tooling.

Vanbronkhorst, J.↗
NASA NTRS1992

Analytical formulation of optimum rotor interdisciplinary design with a three-dimensional wake

An analytical formulation of optimum rotor interdisciplinary design is presented. A finite-state aeroelastic rotor model, coupling generalized dynamic wake with blade finite elements, is applied to perform the optimum rotor blade design for improved aerodynamic performance and vehicle vibration, while a feasible direction nonlinear optimizer, CONMIN, provides the optimization algorithm. The approach features a systematic rotor aeroelastic model which offers an efficient analytical tool, and retains necessary aerodynamic and blade dynamic building blocks for a sufficient rotor dynamic response analysis. The formulation is well suited for an efficient design sensitivity computation without resorting to finite difference, and thus provides a practical design tool. The results show improved rotor aerodynamic performance and reduced hub vibratory loads for the optimized blade as compared to the advanced rotor of reference design.

He, Chengjian↗
NASA NTRS2004

Rotor Design Options for Improving XV-15 Whirl-Flutter Stability Margins

Rotor design changes intended to improve tiltrotor whirl-flutter stability margins were analyzed. A baseline analytical model of the XV-15 was established, and then a thinner, composite wing was designed to be representative of a high-speed tiltrotor. The rotor blade design was modified to increase the stability speed margin for the thin-wing design. Small rearward offsets of the aerodynamic-center locus with respect to the blade elastic axis created large increases in the stability boundary. The effect was strongest for offsets at the outboard part of the blade, where an offset of the aerodynamic center by 10% of tip chord improved the stability margin by over 100 knots. Forward offsets of the blade center of gravity had similar but less pronounced effects. Equivalent results were seen for swept-tip blades. Appropriate combinations of sweep and pitch stiffness completely eliminated whirl flutter within the speed range examined; alternatively, they allowed large increases in pitch-flap coupling (delta-three) for a given stability margin. A limited investigation of the rotor loads in helicopter and airplane configuration showed only minor increases in loads.

Acree, C. W., Jr.↗
NASA NTRS2003

Effects of Blade Sweep on V-22 Whirl Flutter and Loads

A CAMRAD II model of the V-22 Osprey tiltrotor was constructed for the purpose of analyzing the effects of blade design changes on whirl flutter. The model incorporated a dual load-path griplyoke assembly, a swashplate coupled to the transmission, and a drive train. A multiple-trailer free wake was used for loads calculations. The effects of rotor design changes on whirl-mode stability were calculated for swept blades and offset tip masses. A rotor with swept tips and inboard tuning masses was examined in detail to reveal the mechanisms by which these design changes affect stability and loads. Certain combinations of design features greatly increased whirl-mode stability, with (at worst) moderate increases to loads.

Acree, C. W., Jr.↗
NASA NTRS2004

Effects of V-22 Blade Modifications on Whirl Flutter and Loads

A CAMRAD II model of the V-22 Osprey tiltrotor was constructed for the purpose of analyzing the effects of blade design changes on whirl flutter. The model incorporated a dual load-path grip/yoke assembly, a swashplate coupled to the transmission case, and a drive train. A multiple-trailer free wake was used for loads calculations. The effects of rotor design changes on whirl-mode stability were calculated for swept blades and offset tip masses. A rotor with swept tips and inboard tuning masses was examined in detail to reveal the mechanisms by which these design changes affect stability and loads. Certain combinations of design features greatly increased whirl-mode stability, with (at worst) moderate increases to loads.

Acree, C. W., Jr.↗
NASA NTRS2005

Effects of Swept Tips on V-22 Whirl Flutter and Loads

A CAMRAD II model of the V-22 Osprey tiltrotor was constructed for the purpose of analyzing the effects of blade design changes on whirl flutter. The model incorporated a dual load-path grip/yoke assembly, a swashplate coupled to the transmission case, and a drive train. A multiple-trailer free wake was used for loads calculations. The effects of rotor design changes on whirl-mode stability were calculated for swept blades and offset tip masses. A rotor with swept tips and inboard tuning masses was examined in detail to reveal the mechanisms by which these design changes affect stability and loads. Certain combinations of design features greatly increased whirl-mode stability, with (at worst) moderate increases to loads.

Acree, C. W., Jr.↗
NASA NTRS1984

Advanced Gas Turbine (AGT) powertrain system development for automotive applications

Rotor dynamic instability investigations were conducted. Forward ball bearing hydraulic mount configurations were tested with little effect. Trial assembly of S/N 002 ceramic engine was initiated. Impeller design activities were completed on the straight line element (SLE) blade definition to address near-net-shape powder metal die forging. Performance characteristics of the Baseline Test 2A impeller were closely preserved. The modified blading design has been released for tooling procurement. Developmental testing of the diffusion flame combustor (DFC) for initial use in the S/N 002 2100 F ceramic structures engine was completed. A natural gas slave preheater was designed and fabricated. Preliminary regenerator static seal rig testing showed a significant reduction in leakage and sensitivity to stack height. Ceramic screening tests were completed and two complete sets of ceramic static structures were qualified for engine testing. Efforts on rotor dynamics development to resolve subsynchronous motion were continued.

Source record↗
NASA NTRS1997

Euler/Navier-Stokes Solvers Applied to Ducted Fan Configurations

Due to noise considerations, ultra high bypass ducted fans have become a more viable design. These ducted fans typically consist of a rotor stage containing a wide chord fan and a stator stage. One of the concerns for this design is the classical flutter that keeps occurring in various unducted fan blade designs. These flutter are catastrophic and are to be avoided in the flight envelope of the engine. Some numerical investigations by Williams, Cho and Dalton, have suggested that a duct around a propeller makes it more unstable. This needs to be further investigated. In order to design an engine to safely perform a set of desired tasks, accurate information of the stresses on the blade during the entire cycle of blade motion is required. This requirement in turn demands that accurate knowledge of steady and unsteady blade loading be available. Aerodynamic solvers based on unsteady three-dimensional analysis will provide accurate and fast solutions and are best suited for aeroelastic analysis. The Euler solvers capture significant physics of the flowfield and are reasonably fast. An aerodynamic solver Ref. based on Euler equations had been developed under a separate grant from NASA Lewis in the past. Under the current grant, this solver has been modified to calculate the aeroelastic characteristics of unducted and ducted rotors. Even though, the aeroelastic solver based on three-dimensional Euler equations is computationally efficient, it is still very expensive to investigate the effects of multiple stages on the aeroelastic characteristics. In order to investigate the effects of multiple stages, a two-dimensional multi stage aeroelastic solver was also developed under this task, in collaboration with Dr. T. S. R. Reddy of the University of Toledo. Both of these solvers were applied to several test cases and validated against experimental data, where available.

Keith, Theo G., Jr.↗
NASA NTRS1976

PMR polyimide/graphite fiber composite fan blades

Ultrahigh speed fan blades, designed in accordance with the requirements of an ultrahigh tip speed blade axial flow compressor, were fabricated from a high strength graphite fiber tow and a PMR polyimide resin. The PMR matrix was prepared by combining three monomeric reactants in methyl alcohol, and the solution was applied directly to the reinforcing fiber for subsequent in situ polymerization. Some of the molded blades were completely finished by secondary bonding of root pressure pads and an electroformed nickel leading edge sheath prior to final machining. The results of the spin testing of nine PMR fan blades are given. Prior to blade fabrication, heat resin tensile properties of the PMR resin were examined at four formulated molecular weight levels. Additionally, three formulated molecular weight levels were investigated in composite form with both a high modulus and a high strength fiber, both as-molded and postcured, in room temperature and 232 C transverse tensile, flexure and short beam shear. Mixed fiber orientation panels simulating potential blade constructions were also evaluated. Flexure tests, short beam shear tests, and tensile tests were conducted on these angle-plied laminates.

Cavano, P. J.↗
NASA NTRS1951

Some effects of blade trailing-edge thickness on performance of a single-stage axial-flow compressor

A set of modified NACA 65-series blower blades designed for axial inlet velocity, high inlet Mach number, and high blade loading was investigated for trailing-edge thickness of 0.015, 0.030, and 0.045 inch to determine the effect of trailing-edge thickness on single-stage axial-flow-compressor performance. Trailing-edge thickness effects were small except at the highest tip speed investigated (915 ft/sec). Trailing-edge thickness up to 30 percent of maximum blade thickness were used without sacrifice of performance of NACA 65-series blades.

AIRFOILS - THICKNESS↗
NASA NTRS2010

SSME Investment in Turbomachinery Inducer Impeller Design Tools and Methodology

Within the rocket engine industry, SSME turbomachines are the de facto standards of success with regard to meeting aggressive performance requirements under challenging operational environments. Over the Shuttle era, SSME has invested heavily in our national inducer impeller design infrastructure. While both low and high pressure turbopump failures/anomaly resolution efforts spurred some of these investments, the SSME program was a major benefactor of key areas of turbomachinery inducer-impeller research outside of flight manifest pressures. Over the past several decades, key turbopump internal environments have been interrogated via highly instrumented hot-fire and cold-flow testing. Likewise, SSME has sponsored the advancement of time accurate and cavitating inducer impeller computation fluid dynamics (CFD) tools. These investments together have led to a better understanding of the complex internal flow fields within aggressive high performing inducers and impellers. New design tools and methodologies have evolved which intend to provide confident blade designs which strike an appropriate balance between performance and self induced load management.

Zoladz, Thomas↗
NASA NTRS1984

Helicopter rotor blade aerodynamic optimization by mathematical programming

Formal mathematical programing was applied to the aerodynamic rotor blade design process. The approach is to couple hover and forward flight analysis programs with the general-purpose optimization program CONMIN to determine the blade taper ratio, percent taper, twist distribution, and solidity which minimize the horsepower required at hover while meeting constraints on forward flight performance. Designs obtained using this approach for the blade of a representative Army helicopter compare well with those obtained using a conventional approach involving personnel-intensive parametric studies. Results from the present method can be obtained in 2 days as compared to 5 weeks required by the conventional procedure. Also the systematic manipulation of the design variables by the optimization procedure minimizes the need for the researcher to have a vast body of past experience and data in determining the influence of a design change on the performance.

Walsh, J. L.↗
NASA NTRS1994

Aeroelastic Stability of Modern Bearingless Rotors: A Parametric Investigation

The University of Maryland Advanced Rotorcraft Code (UMARC) is utilized to study the effects of blade design parameters on the aeroelastic stability of an isolated modern bearingless rotor blade in hover. The McDonnell Douglas Advanced Rotor Technology (MDART) Rotor is the baseline rotor investigated. Results indicate that kinematic pitch-lag coupling introduced through the control system geometry and the damping levels of the shear lag dampers strongly affect the hover inplane damping of the baseline rotor blade. Hub precone, pitchcase chordwise stiffness, and blade fundamental torsion frequency have small to moderate influence on the inplane damping, while blade pre-twist and placements of blade fundamental flapwise and chord-wise frequencies have negligible effects. A damperless configuration with a leading edge pitch-link, 15 deg of pitch-link cant angle, and reduced pitch-link stiffness is shown to be stable with an inplane damping level in excess of 2.7 percent critical at the full hover tip speed.

Nguyen, Khanh Q.↗
NASA NTRS1996

BladeCAD: An Interactive Geometric Design Tool for Turbomachinery Blades

A new metthodology for interactive design of turbomachinery blades is presented. Software implementation of the meth- ods provides a user interface that is intuitive to aero-designers while operating with standardized geometric forms. The primary contribution is that blade sections may be defined with respect to general surfaces of revolution which may be defined to represent the path of fluid flow through the turbomachine. The completed blade design is represented as a non-uniform rational B-spline (NURBS) surface and is written to a standard IGES file which is portable to most design, analysis, and manufacturing applications.

Miller, Perry L., IV↗