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Calculation is taken from pressure or velocity distribution along blade surface. Blade sections have either one or two /tandem/ blade segments. Program is recommended primarily for use with ideal flow pressure distributions where forces in drag direction are neglected.
Stripper incorporates a blade which is electrically heated to a controlled temperature. Heated blade softens and strips insulation from cable while paper ribbon removes insulation material and keeps blade clean for next operation.
Machining of slot in turbine blades of axial flow turbines to provide flow path between pressure and suction surfaces is discussed. Slot configuration and improvements in blade performance are described. Diagram of blade slot to show geometry of modification is included.
The results of an analytical study to determine the effect of changes in the amount, velocity, injection location, injection angle, and temperature of coolant flow on blade row performance are presented. The results show that the change in output of a cooled turbine blade row relative to the specific output of the uncooled blade row can be positive, negative, or zero. Comparisons between the analytical results and experimental results for four different cases of coolant discharge, all at a coolant temperature ratio of unity, show good agreement for three cases and rather poor agreement for the other. To further test the validity of the method, more experimental data is needed, particularly at different coolant temperature ratios.
Blade slap encountered in rotary wings and its effect on helicopter performance are reported. The results of various individual flight tests are presented and, where possible, correlated with one another. Observations from the subjective evaluation of blade slap are included, together with a modified form of the blade slap factor (BSF) which can be used as a design criteria.
A bidirectional jet flap device called the variable deflection thruster (VDT) has been investigated for possible application to helicopter rotors. This investigation included the development and testing of a fluidic lift control system for the VDT-blade model making use of the test result that VDT-blade lift can be sensed from the differential pressure at midchord. This study constitutes a long-range program to develop blown control techniques for stabilizing the higher harmonic modes of helicopter rotors. Wind tunnel tests were conducted using a three-sectioned, two-dimensional VDT-blade model having individually controlled VDT jet flaps in each section. Steady-state tests were conducted without the fluidic lift controller (open loop) for both full-span blowing and for the model center section blowing only. Steady-state tests were conducted with the center section blowing only using the fluidic lift controller (close-loop) to control the lift on the model center section. Dynamic tests were conducted using the complete model with the VDT jet in the model center section oscillating at various frequencies and also using the model center section alone on a single endplate to obtain finite-aspect-ratio effects. Fair agreement was obtained between theory and experimental results.
A conventional rotor and stator, two dual-airfoil tandem rotors, and one dual-airfoil tandem stator were designed. The two tandem rotors were each designed with different percentages of the overall lift produced by the front airfoil. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. Rotor inlet hub/tip ratio was 0.8. Design values of rotor tip velocity and stage pressure ratio were 757 ft/sec and 1.30, respectively.
A conceptual optical method and some test results are presented for measuring the surface temperature distribution on one of the rotating turbine blades with a surface resolution of 0.05 cm spot diameter at a tip speed of 400 m/sec. The blade is scanned line-by-line by a fixed optical system. During each line-scan, the detector analog output signal is converted to two hundred consecutive digital values that are temporarily stored in a high speed buffer memory, and then transferred at a slower rate to a computer for processing. The signal-to-noise ratio of the silicon avalanche detector is large enough to obtain an accuracy of one percent at 1050 K blade temperature. By averaging 25 scans of the same line the same accuracy can be obtained at 900 K.
A traverse mechanism which allows the measurement of the three dimensional boundary layers on a helicopter rotor blade has been built and tested on a full scale rotor to full scale conditions producing centrifugal accelerations in excess of 400 g and Mach numbers of 0.6 and above. Boundary layer velocity profiles have been measured over a range of rotor speeds and blade collective pitch angles. A pressure scanning switch and transducer were also tested on the full scale rotor and found to be insensitive to centrifugal effects within the normal main rotor operating range. The demonstration of the capability to measure boundary layer behavior on helicopter rotor blades represents the first step toward obtaining, in the rotating system, data of a quality comparable to that already existing for flows in the fixed system.
The flow characteristics of several tandem bladed compressor stators were analytically evaluated over a range of inlet incidence angles. The ratios of rear-segment to front-segment chord and camber were varied. Results were also compared to the analytical performance of a reference solid blade section. All tandem blade sections exhibited lower calculated losses than the solid stator. But no one geometric configuration exhibited clearly superior characteristics. The front segment accepts the major effect of overall incidence angle change. Rear- to front-segment camber ratios of 4 and greater appeared to be limited by boundary-layer separation from the pressure surface of the rear segment.
A conceptual optical method and some test results are presented for measuring the surface temperature distribution on one of the rotating turbine blades with a surface resolution of 0.05 cm spot diameter at a tip speed of 400 m/sec. The blade is scanned line-by-line by a fixed optical system. During each line-scan, the detector analog output signal is converted to 200 consecutive digital values that are temporarily stored in a high-speed buffer memory and then transferred at a slower rate to a computer for processing. The signal-to-noise ratio of the silicon avalanche detector is large enough to obtain an accuracy of 1% at 1050 K blade temperature. By averaging 25 scans of the same line the same accuracy can be obtained at 900 K.
In this study the stability boundaries of a hingeless helicopter blade are studied using a system of coupled flap-lag-pitch equations of motion. Divergence and flutter boundaries for the linearized system of equations are presented. Flap-lag, flap-pitch and coupled flap-lag-pitch instabilities in hovering flight are studied. The effect of the torsional degree of freedom on the flap-lag type of instability is investigated. Similarly the effect of lag on the flap pitch type of instability is also considered. Results illustrating these effects, together with the effect of various important blade parameters on blade stability are presented.
An axial flow compressor stage, having single-airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor had an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were: (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of tandem-airfoil blading designed for the same vector diagrams; and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.
The complexities of turbine engine blade vibration are compounded by blade geometry, temperature gradients, and rotational speeds. Experience indicates that dynamics analysis using the finite element approach provides an effective means for predicting vibration characteristics of compressor and turbine blades whose geometry may be irregular, have curved boundaries, and be subjected to high temperatures and speeds. The NASTRAN program was chosen to help analyze the dynamics of normal modes, rotational stiffening and thermal effects on the normal modes, and forced responses. The program has produced reasonable success. This paper presents the analytical procedures and the NASTRAN results, in comparison with a conventional beam element program and laboratory data.
An axial flow compressor stage, having tandem airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor has an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of single-airfoil blading designed for the same vector diagrams and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design.
A theoretical analysis was conducted of the dynamic behavior of micron size particles entrained in gas flow on the two-dimensional blade-to-blade surface of a circular stationary cascade of turbine stator blades. The particle velocity lag and angular deviation relative to the gas was determined as a function of particle diameter and mass density. Particles size and density were varied over ranges selected to correspond to typical laser-Doppler velocimeter (LDV) flow field mapping applications. It was found that velocity lag and angular deviation increased whenever particle size or mass density increased, and that particle tracking was more sensitive to a change in particle diameter than to a change in mass density. Results indicated that LDV applications employing 1 gm/cc tracer particles with diameters greater than approximately 1 micron, or 0.5 micron diameter particles with mass densities greater than 4 gm/cc would experience velocity and angular deviations generally greater than 2 percent and 1 degree, respectively.
Two holographic interblade row flow visualization systems were designed to determine the three-dimensional shock patterns and velocity distributions within the rotating blade row of a transonic fan rotor, utilizing the techniques of pulsed laser transmission holography. Both single- and double-exposure bright field holograms and dark field scattered-light holograms were successfully recorded. Two plastic windows were installed in the rotor tip casing and outer casing forward of the rotor to view the rotor blade passage. The viewing angle allowed detailed investigation of the leading edge shocks and shocks in the midspan damper area; limited details of the trailing edge shocks also were visible. A technique was devised for interpreting the reconstructed holograms by constructing three dimensional models that allowed identification of the major shock systems. The models compared favorably with theoretical predictions and results of the overall and blade element data. Most of the holograms were made using the rapid double-pulse technique.
The effect of incidence angle on the aerodynamic performance of an uncooled core turbine rotor blade was investigated experimentally in a two-dimensional cascade. The cascade test covered a range of incidence angles from minus 15 deg to 15 deg in 5-degree increments and a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95. The principal measurements were blade-surface static pressures and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the investigation include blade-surface velocity distribution and overall performance in terms of weight flow and loss for the range of incidence angles and exit velocity ratios investigated. The measured losses are also compared with two common methods of predicting incidence loss.