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The methodology for fabrication of wet filament wound spar shell fan blades is presented. All principal structural elements were filament wound, assembled, formed, bonded and co-cured in a female mold. A pair of blades were fabricated as one integral unit and parted into two after curing.
The results are presented of a one-dimensional dynamic digital blade row compressor model study of a J85-13 engine operating with uniform and with circumferentially distorted inlet flow. Details of the geometry and the derived blade row characteristics used to simulate the clean inlet performance are given. A stability criterion based upon the self developing unsteady internal flows near surge provided an accurate determination of the clean inlet surge line. The basic model was modified to include an arbitrary extent multi-sector parallel compressor configuration for investigating 180 deg 1/rev total pressure, total temperature, and combined total pressure and total temperature distortions. The combined distortions included opposed, coincident, and 90 deg overlapped patterns. The predicted losses in surge pressure ratio matched the measured data trends at all speeds and gave accurate predictions at high corrected speeds where the slope of the speed lines approached the vertical.
A mathematical model is developed to evaluate a number of physical effects in turbomachine blade tip contact on the basis of a broad brush approach to provide insights without undue complexity. The analysis is restricted to the case of a shoe encountering a moving blade row and pressing against it with a fixed time-average force. Relative displacement components in the contact region are identified and determined. Two key equations are derived, where it is shown that wear rate plays a strong role in determining the traversal speed of the contact patch. The formula for estimating the size of the contact patch shows that the higher the traversal speed is, the wider is the contact patch.
This paper describes a generalized dynamic model which has been developed for use in compression component aerodynamic stability studies. The model is a one-dimensional, pitch-line, blade row, lumped volume system. Arbitrary placement of blade free volumes upstream, within, and downstream of the compression component as well as the removal of bleed flow from the exit of any rotor or stator are model options. The model has been applied to a two-stage fan and an eight-stage compressor. The clean inlet pressure ratio/flow maps and the surge line have been reproduced, thereby validating the capability of the dynamic model to reproduce the steady-flow characteristics of the compression component. A method for determining the onset of an aerodynamic instability which is associated with surge is described. Sinusoidally time-varying inlet and exit boundary conditions have been applied to the eight stage compressor as examples of the manner in which this model may be used for stability studies.
A current NASA-sponsored contract will develop attachment technology for hybrid turbine rotors having dovetailed ceramic blades and a wrought metal disk. The approach and sequence of activities are described in detail. The program scope includes design, fabrication, and spin testing of hot-pressed silicon nitride blades for operation at rotor inlet temperatures of 2192 F (1200 C) or greater.
Various existing unsteady aerodynamic strip theories which have been developed in the past for both fixed and rotary wing aeroelastic analyses are modified in the paper so as to make them applicable to the coupled flap-lag-torsional aeroelastic problem of a rotor blade in hover. These corrections are primarily due to constant angle of attack, constant inflow and variable free stream velocity due to lead-lag motion. Next, the modified strip theories are incorporated in a coupled flap-lag-torsional aeroelastic analysis of the rotor blade in hover and the sensitivity of the aeroelastic stability boundaries to the aerodynamic assumptions is examined.
The generation of blade passing frequency noise due to rotor wakes interaction with the downstream stator is examined using a recent model for the unsteady aerodynamic forces developed by Goldstein and Atassi. The effects of blade geometry and loading are investigated. It is shown that at low reduced frequencies loading significantly reduces the noise level while at high frequencies loading substantially increases the sound level radiated. It is also shown that for the same loading a higher angle of attack generates more noise particularly at higher reduced frequencies. The noise level shows a quasi-periodic pattern with pronounced minima and maxima as the frequency increases. These conclusions are valid for different wake models.
The differential equations of motion for the lateral and torsional deformations of a nonlinearly twisted rotor blade in steady flight conditions together with those additional aeroelastic features germane to composite bearingless rotors are derived. The differential equations are formulated in terms of uncoupled (zero pitch and twist) vibratory modes with exact coupling effects due to finite, time variable blade pitch and, to second order, twist. Also presented are derivations of the fully coupled inertia and aerodynamic load distributions, automatic pitch change coupling effects, structural redundancy characteristics of the composite bearingless rotor flexbeam - torque tube system in bending and torsion, and a description of the linearized equations appropriate for eigensolution analyses. Three appendixes are included presenting material appropriate to the digital computer program implementation of the analysis, program G400.
The nonlinear curvature expressions for a twisted rotor blade or a beam undergoing transverse bending in two planes, torsion, and extension were developed. The curvature expressions were obtained using simple geometric considerations. The expressions were first developed in a general manner using the geometrical nonlinear theory of elasticity. These general nonlinear expressions were then systematically reduced to four levels of approximation by imposing various simplifying assumptions, and in each of these levels the second degree nonlinear expressions were given. The assumptions were carefully stated and their implications with respect to the nonlinear theory of elasticity as applied to beams were pointed out. The transformation matrices between the deformed and undeformed blade-fixed coordinates, which were needed in the development of the curvature expressions, were also given for three of the levels of approximation. The present curvature expressions and transformation matrices were compared with corresponding expressions existing in the literature.
The effects of several characteristics of helicopter blade slap upon human annoyance are examined. Blade slap noise was simulated by using continuous and impulsive noises characterized by five parameters: The number of sine waves in a single impulse; the frequency of the sine waves; the impulse repetition frequency; the sound pressure level (SPL) of the continuous noise; and the idealized crest factor of the impulses. Ten second samples of noise were synthesized with each of the five parameters at representative levels. The annoyance of each noise was judged by 40 human subjects. Analysis of the subjective data indicated that each of the five parameters had a statistically significant effect upon the annoyance judgments. The impulse crest factor and SPL of the continuous noise had very strong positive relationships with annoyance. The other parameters had smaller, but still significant, effects upon the annoyance judgments.
The overall and blade-element performance of a low pressure ratio, low tip speed fan stage at design speed is presented for tow off-design rotor blade angle settings. The rotor design tip speed is 243.8 m/sec and weight flow per unit annulus area is 175.8 kg/sec. Design weight flow and pressure ratio are 29.9 kg/sec and 1.151, respectively.
A two dimensional finite difference numerical technique is presented to determine the temperature distribution of an internal cooled blade of radial turbine guide vanes. A simple convection cooling is assumed inside the guide vane blade. Such cooling has relatively small cooling effectiveness at the leading edge and at the trailing edge. Heat transfer augmentation in these critical areas may be achieved by using impingement jets and film cooling. A computer program is written in FORTRAN IV for IBM 370/165 computer.
NASA developed stability and frequency response analysis techniques were applied to a dynamic blade row compression component stability model to provide a more economic approach to surge line and frequency response determination than that provided by time-dependent methods. This blade row model was linearized and the Jacobian matrix was formed. The clean-inlet-flow stability characteristics of the compressors of two J85-13 engines were predicted by applying the alternate Routh-Hurwitz stability criterion to the Jacobian matrix. The predicted surge line agreed with the clean-inlet-flow surge line predicted by the time-dependent method to a high degree except for one engine at 94% corrected speed. No satisfactory explanation of this discrepancy was found. The frequency response of the linearized system was determined by evaluating its Laplace transfer function. The results of the linearized-frequency-response analysis agree with the time-dependent results when the time-dependent inlet total-pressure and exit-flow function amplitude boundary conditions are less than 1 percent and 3 percent, respectively. The stability analysis technique was extended to a two-sector parallel compressor model with and without interstage crossflow and predictions were carried out for total-pressure distortion extents of 180 deg, 90 deg, 60 deg, and 30 deg.
The results of dynamic analyses, weight and balance tests, static stiffness tests, and structural vibration tests on the 60-foot-long metal blades for the ERDA-NASA 100-kW wind turbine are presented. The metal blades are shown to be free from structural or dynamic resonance at the wind turbine design speed. Aeroelastic instabilities are unlikely to occur within the normal operating range of the wind turbine.
The performance of the first stage of a transonic, multistage compressor was mapped over a range of inlet-guide-vane and stator-blade settings. Both stall-free and deep-stall performance data were obtained. For the settings tested, as stall was encountered and flow was further reduced, a relatively sharp drop in pressure ratio occurred and was followed by a continuing but more gradual reduction in pressure ratio with reduced flow. The position of the stall line on the map of pressure ratio against equivalent weight flow was essentially unaffected over the range of inlet-guide-vane and stator-blade settings.
A research investigation was conducted to assess the potential of the Tip Air Mass Injection system in reducing the noise output during blade vortex interaction in descending low speed flight. In general it was concluded that the noise output due to blade vortex interaction can be reduced by 4 to 6 db with an equivalent power expenditure of approximately 14 percent of installed power.
The dynamic stability of the flap/lag/torsion motion of a uniform, cantilever rotor blade in hover is calculated. The influence of blade collective pitch, lag frequency, torsional flexibility, structural coupling, and precone angle on the stability is examined. Good agreement is found with the results of an independent analytical investigation.
Pressure transducers were installed on the blades and vanes of QF-1B, a transonic-tip-speed fan from the NASA Quiet Fan Program which was tested on the outdoor quiet fan test facility at NASA-Lewis. Signals from the transducers and from far field microphones were analyzed to determine sources of nonuniform inflow and noise. The nonuniform inflow was mostly unsteady with roughly equal contributions from atmospheric turbulence and rig interference. The rig interference was largest at the bottom and appeared to be generated by the support structure which was located behind the inlet lip under the fan. Interaction of this inflow distortion was the dominant source of noise at 1, 2, and 3 times blade passing frequency (BPF) at 60, 70, and 80 percent of design speed. At 90 percent speed, noise at BPF was dominated by the steady rotor field. A broadband spectrum peak centered at about 2.2 times BPF was identified as rotor/stator interaction stemming from a high frequency rotor exit flow component. The remaining broadband energy from 0.3 to 3.5 times BPF was attributed to the better known type of rotor/stator interaction associated with rotor wake turbulence.