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At least 325 records · Page 18

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NASA NTRS1959

Evaluation of Transpiration-Cooled Turbine Blades with Shells of "Poroloy" Wire Cloth

An experimental investigation was made to evaluate the durability and permeability of a group of transpiration-cooled, strut-supported turbine blades. The porous shells were formed from a woven-wire material. The blades were fabricated by a contractor for the Bureau of Aeronautics. The results of permeability tests indicated that the shell material exhibited large random variations in local permeability, which result in excessive coolant flows and very nonuniform cooling. For this reason no heat-transfer evaluations were made because any results would have been inconclusive. Four blades were investigated for structural soundness in a turbo-jet engine operating at a turbine-inlet temperature of approximately 1670 deg F and a turbine tip speed of approximately 1305 feet per second. The maximum temperature of the porous-shell material was approximately 1050 deg F. Inspection of the first two blades after 10 minutes of engine operation revealed that the tips of both of the blades had failed. For the second pair of blades, an improved tip cap was provided by the use of built-up weld extending from strut tip to shell. One of these blades was then operated for 33 hours without failure, and was found to be in good condition at the end of this time. The second blade of this second pair failed within the first 10 minutes of operation because of a poor bond between shell and strut lands.

Richards, Hadley T.↗

NASA NTRS1999

Dual-Laser Probe for Measuring Blade-Tip Clearance Tested

A dual-probe, integrated fiber-optic laser system for measuring blade tip clearance in rotating turbomachinery was developed cooperatively with Integrated Fiber Optic Systems, Inc., and tested at the NASA Lewis Research Center. Because the probes are nearly flush with the casing inner lining, there is minimal flow disturbance. The two probes are closely spaced in a circumferential plane and are slanted at an angle relative to each other so that the time it takes the blade tip to traverse the space between the two beams varies with the tip radius, making it possible to determine the tip clearance at rotor operating conditions. The tip clearance can be obtained for all the blades in a rotor with a single system, provided there are no synchronous vibrations present at a particular operating condition. The two probes for the laser system were installed in two holders. For one, the included angle between the probes was 20 degrees, and for the other, it was 40 degrees. The two configurations were calibrated in a vacuum spin-rig facility that can reproduce realistic blade-tip speeds. A specially designed, nondeflecting rotor was used to calibrate the probes. This rotor consisted of a tapered titanium bar with three teeth on each end, the thickness of which was representative of a typical blade-tip thickness. The bar and measuring teeth were in a horizontal plane in the spin rig, and the axis of rotation was vertical. To set the clearances, we used a remotely controlled optical stage that could be traversed vertically. The 40 probe system also was used to measure the tip clearance during an engine fan prototype test in a wind tunnel. Using the spin-rig calibrations, we estimated the accuracy of the tip-clearance measurement in the test to be 0.005 in. This program is ongoing.

Dhadwal, Harbans S.↗

NASA NTRS1987

Noise of a model high speed counterrotation propeller at simulated takeoff/approach conditions (F7/A7)

A high speed advanced counterrotation propeller, was tested in the NASA-Lewis 9 x 15 foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Acoustic measurements were taken with fixed floor microphones, an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could take both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish the propeller operating conditions. The propeller was run over a range of blade setting angles from 36.4/36.5 to 41.1/39.4 deg, tip speeds from 165 to 259 m/sec, rotor spacings from 1.56 to 3.63 based on forward rotor tip chord to aerodynamic separation, and angles of attack to + or - 16 deg. First order rotor alone tones showed highest directivity levels near the propeller plane, while interaction tone showed high levels throughout sideline directivity, especially toward the propeller rotation axis. Interaction tone levels were sensitive to propeller row spacing while rotor alone tones showed little spacing effect. There is a decreased noise level associated with higher propeller blade numbers for the same overall propeller thrust.

Woodward, Richard P.↗

NASA NTRS1987

Noise of a model high speed counterrotation propeller at simulated takeoff/approach conditions (F7/A7)

Woodward, Richard P.↗

NASA NTRS1983

NASA Low-Speed Centrifugal Compressor for Fundamental Research

A centrifugal compressor facility being built by the NASA Lewis Research Center is described; its purpose is to obtain benchmark experimental data for internal flow code verification and modeling. The facility will be heavily instrumented with standard pressure and temperature probes and have provisions for flow visualization and laser Doppler velocimetry. The facility will accommodate rotational speeds to 2400 rpm and will be rated at pressures to 1.25 atm. The initial compressor stage for testing is geometrically and dynamically representative of modern high-performance stages with the exception of Mach number levels. Design exit tip speed for the initial stage is 500 ft/sec with a pressure ratio of 1.17. The rotor exit backsweep is 55 deg from radial.

Wood, J. R.↗

NASA NTRS1983

NASA low-speed centrifugal compressor for fundamental research

A new centrifugal compressor facility being built by the NASA Lewis Research Center is described; its purpose is to obtain 'benchmark' experimental data for internal flow code verification and modeling. The facility will be heavily instrumented with standard pressure and temperature probes and have provisions for flow visualization and laser Doppler velocimetry. The facility will accommodate rotational speeds to 2400 rpm and will be rated at pressures to 1.25 atm. The initial compressor stage for testing is geometrically and dynamically representative of modern high-performance stages with the exception of Mach number levels. Design exit tip speed for the initial stage is 500 ft/sec with a pressure ratio of 1.17. The rotor exit backsweep is 55 deg from radial. The facility is expected to be operational in the first half of 1985.

Wood, J. R.↗

NASA NTRS1985

Large-scale Advanced Propfan (LAP) performance, acoustic and weight estimation, January, 1984

In comparison to turbo-prop applications, the Prop-Fan is designed to operate in a significantly higher range of aircraft flight speeds. Two concerns arise regarding operation at very high speeds: aerodynamic performance and noise generation. This data package covers both topics over a broad range of operating conditions for the eight (8) bladed SR-7L Prop-Fan. Operating conditions covered are: Flight Mach Number 0 - 0.85; blade tip speed 600-800 ft/sec; and cruise power loading 20-40 SHP/D2. Prop-Fan weight and weight scaling estimates are also included.

Parzych, D.↗

NASA NTRS1987

Propfan test assessment propfan propulsion system static test report

The propfan test assessment (PTA) propulsion system successfully completed over 50 hours of extensive static ground tests, including a 36 hour endurance test. All major systems performed as expected, verifying that the large-scale 2.74 m diameter propfan, engine, gearbox, controls, subsystems, and flight instrumentation will be satisfactory with minor modifications for the upcoming PTA flight tests on the GII aircraft in early 1987. A test envelope was established for static ground operation to maintain propfan blade stresses within limits for propfan rotational speeds up to 105 percent and power levels up to 3880 kW. Transient tests verified stable, predictable response of engine power and propfan speed controls. Installed engine TSFC was better than expected, probably due to the excellent inlet performance coupled with the supercharging effect of the propfan. Near- and far-field noise spectra contained three dominant components, which were dependent on power, tip speed, and direction. The components were propfan blade tones, propfan random noise, and compressor/propfan interaction noise. No significant turbine noise or combustion noise was evident.

Orourke, D. M.↗

NASA NTRS1989

Propfan Test Assessment (PTA)

The objectives of the Propfan Test Assessment (PTA) Program were to validate in flight the structural integrity of large-scale propfan blades and to measure noise characteristics of the propfan in both near and far fields. All program objectives were met or exceeded, on schedule and under budget. A Gulfstream Aerospace Corporation GII aircraft was modified to provide a testbed for the 2.74m (9 ft) diameter Hamilton Standard SR-7 propfan which was driven by a 4475 kw (600 shp) turboshaft engine mounted on the left-hand wing of the aircraft. Flight research tests were performed for 20 combinations of speed and altitude within a flight envelope that extended to Mach numbers of 0.85 and altitudes of 12,192m (40,000 ft). Propfan blade stress, near-field noise on aircraft surfaces, and cabin noise were recorded. Primary variables were propfan power and tip speed, and the nacelle tilt angle. Extensive low altitude far-field noise tests were made to measure flyover and sideline noise and the lateral attenuation of noise. In coopertion with the FAA, tests were also made of flyover noise for the aircraft at 6100m (20,000 ft) and 10,668m (35,000 ft). A final series of tests were flown to evaluate an advanced cabin wall noise treatment that was produced under a separate program by NASA-Langley Research Center.

Little, B. H.↗

NASA NTRS1987

An experimental investigation of an advanced turboprop installation on a swept wing at subsonic and transonic speeds

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of a turboprop-nacelle installation on the pressure distributions over a swept, supercritical wing. The tests were conducted at Mach numbers from 0.20 to 0.80, at angles of attack from 0 to 5 degrees, nacelle nozzle pressure ratios from 1.0 to 1.6, and at propeller tip speeds from 700 to 800 ft/sec. The results of this study indicate that the turboprop nacelle interference, with and without power, on a swept wing is greater on the inboard wing panel than on the outboard wing panel. The over-the-wing nacelle installation with the propeller upwash on the inboard panel had flow separation problems at a Mach number of 0.80. No severe flow separation problems appear to exist for either propeller rotation direction for the under-the-wing nacelle installation. The local flow disturbances caused by the under-the-wing nacelle installation were in general less severe than for the over-the-wing nacelle installation.

Carlson, John R.↗

NASA NTRS1973

Holographic studies of shock waves within transonic fan rotors

Pulsed laser holographic interferometry has been applied to the detection of shock patterns in the outer span regions of high tip speed transonic rotors. The first holographic approach used ruby laser light reflected from a portion of the centerbody just ahead of the rotor. These holograms showed the bow wave patterns upstream of the rotor and the shock patterns just inside the blade row near the tip. Much of the region of interest was in the shadow of the blade leading edge and could not be visualized. The second holographic approach, on a different rotor, used light transmitted diagonally across the inlet annulus past the centerbody. This approach gave a more extensive view of the region bounded by the blade leading and trailing edges, by the part span shroud and by the blade tip. These holograms showed the passage shock emanating from the blade leading edge and a moderately strong conical shock originating at the intersection of the part span shroud leading edge and the blade suction surface. Reasonable details of the shock patterns were obtained from holograms which were made without extensive rig modifications.

Benser, W. A.↗

NASA NTRS1974

Holographic studies of shock waves within transonic fan rotors

NASA has funded two separate contracts to apply pulsed laser holographic interferometry to the detection of shock patterns in the outer span regions of high tip speed transonic rotors. The first holographic approach used ruby laser light reflected from a portion of the centerbody just ahead of the rotor. These holograms showed the bow wave patterns upstream of the rotor and the shock patterns just inside the blade row near the tip. The second holographic approach, on a different rotor, used light transmitted diagonally across the inlet annulus past the centerbody. This approach gave a more extensive view of the region bounded by the blade leading and trailing edges, by the part span shroud and by the blade tip. These holograms showed the passage shock emanating from the blade leading edge and a moderately strong conical shock originating at the intersection of the part span shroud leading edge and the blade suction surface.

Benser, W. A.↗

NASA NTRS2012

Influence of Alternative Engine Concepts on LCTR2 Sizing and Mission Profile

The Large Civil Tiltrotor (LCTR) was developed as part of the NASA Heavy Lift Rotorcraft Systems Investigation in order to establish a consistent basis for evaluating the benefits of advanced technology for large tiltrotors. The concept has since evolved into the second-generation LCTR2, designed to carry 90 passengers for 1,000 nm at 300 knots, with vertical takeoff and landing. This paper examines the impact of advanced propulsion system concepts on LCTR2 sizing. Two concepts were studied: an advanced, single-speed engine with a conventional power turbine layout (Advanced Conventional Engine, or ACE), and a variable-speed power turbine engine (VSPT). The ACE is the lighter engine, but requires a multi-speed (shifting) gearbox, whereas the VSPT uses a lighter, fixed-ratio gearbox. The NASA Design and Analysis of Rotorcraft (NDARC) design code was used to study the trades between rotor and engine efficiency and weight. Rotor performance was determined by Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD II), and engine performance was estimated with the Numerical Propulsion System Simulation (NPSS). Design trades for the ACE vs. VSPT are presented in terms of vehicle weight empty for variations in mission altitude and range; the effect of different One Engine Inoperative (OEI) criteria are also examined. Because of its strong effect on gearbox weight and on both rotor and engine efficiency, rotor speed was chosen as the reference design variable for comparing design trades. The two propulsion concepts had nearly identical vehicle weights and mission fuel consumption, and their relative advantages varied little with cruise altitude, mission range, or OEI criteria; high cruise altitude and low cruise tip speed were beneficial for both concepts.

Acree, C. W., Jr.↗

NASA NTRS2012

Quiet High Speed Fan II (QHSF II): Final Report

This report details the aerodynamic, mechanical, structural design and fabrication of a Honey Engines Quiet High Speed Fan II (lower hub/tip ratio and higher specific flow than the Baseline I fan). This fan/nacelle system incorporates features such as advanced forward sweep and an advanced integrated fan/fan exit guide vane design that provides for the following characteristics: (1) Reduced noise at supersonic tip speeds, in comparison to current state-of-the-art fan technology; (2) Improved aeroelastic stability within the anticipated operating envelope; and (3) Aerodynamic performance consistent with current state-of-the-art fan technology. This fan was fabricated by Honeywell and tested in the NASA Glenn 9- by 15-Ft Low Speed Wind Tunnel for aerodynamic, aeromechanical, and acoustic performance.

Kontos, Karen↗

NASA NTRS1979

Low-speed wind-tunnel investigation of a large-scale VTOL lift-fan transport model

An investigation was conducted in the NASA-Ames 40 by 80 Foot Wind Tunnel to determine the aerodynamic characteristics of a large scale, VTOL, lift fan, jet transport model. The model had two lift fans at the forward portion of the fuselage, a lift fan at each wing tip, and two lift/cruise fans at the aft portion of the fuselage. All fans were driven by tip turbines using T-58 gas generators. Results were obtained for several lift fan, exit vane deflections and lift/cruise fan thrust deflections are zero sideslip. Three component longitudinal data are presented at several fan tip speed ratios. A limited amount of six component data were obtained with asymmetric vane settings. All of the data were obtained without a horizontal tail. Downwash angles at a typical tail location are also presented.

Aoyagi, K.↗

NASA NTRS1981

Additional noise data on the SR-3 propeller

The noise generated by supersonic-tip-speed propellers is investigated. An eight bladed propeller was tested in the Lewis 8- by 6-foot wind tunnel with conditions providing data in the subsonic operating region of the propeller. These conditions resulted in a slight reshaping of the curve for blade passing tone as a function of helical tip Mach number as compared with previous results. Directivity curves with an additional transducer position gave an indication of a lobe pattern for this propeller that was not previously observed. The present data at the aft-most position indicate that some reflections, possibly from the test rig support strut, may have affected the data taken previously.

Dittmar, J. H.↗

NASA NTRS1977

Wind tunnel investigation of rotor lift and propulsive force at high speed: Data analysis

The basic test data obtained during the lift-propulsive force limit wind tunnel test conducted on a scale model CH-47b rotor are analyzed. Included are the rotor control positions, blade loads and six components of rotor force and moment, corrected for hub tares. Performance and blade loads are presented as the rotor lift limit is approached at fixed levels of rotor propulsive force coefficients and rotor tip speeds. Performance and blade load trends are documented for fixed levels of rotor lift coefficient as propulsive force is increased to the maximum obtainable by the model rotor. Test data is also included that defines the effect of stall proximity on rotor control power. The basic test data plots are presented in volumes 2 and 3.

Mchugh, F.↗

NASA NTRS1989

Cruise noise of the SR-2 propeller model in a wind tunnel

Noise data on the SR-2 model propeller were taken in the NASA Lewis Research Center 8- by 6-Foot Wind Tunnel. The maximum blade passing tone rises with increasing helical tip Mach number to a peak level at a helical tip Mach number of about 1.05; then it remains the same or decreases at higher helical tip Mach numbers. This behavior, which has been observed with other propeller models, points to the possibility of using higher propeller tip speeds to limit airplane cabin noise while maintaining high flight speed and efficiency. Noise comparisons of the straight-blade SR-2 propeller and the swept-blade SR-7A propeller showed that the tailored sweep of the SR-7A appears to be the cause of both lower peak noise levels and a slower noise increase with increasing helical tip Mach number.

Dittmar, James H.↗