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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1954

Calibration of strain-gage installations in aircraft structures for the measurement of flight loads

A general method has been developed for calibrating strain-gage installations in aircraft structures, which permits the measurement in flight of the shear or lift, the bending moment, and the torque or pitching moment on the principal lifting or control surfaces. Although the stress in structural members may not be a simple function of the three loads of interest, a straightforward procedure is given for numerically combining the outputs of several bridges in such a way that the loads may be obtained. Extensions of the basic procedure by means of electrical combination of the strain-gage bridges are described which permit compromises between strain-gage installation time, availability of recording instruments, and data reduction time. The basic principles of strain-gage calibration procedures are illustrated by reference to the data for two aircraft structures of typical construction, one a straight and the other a swept horizontal stabilizer.

Skopinski, T H↗
NASA NTRS1945

An Investigation of the Ranger V-770-8 Engine Installation for the Edo XOSE-1 Airplane I : Cooling

Engine temperature data and cooling correlating analyses of the engine and oil cooler are presented in connection with an investigation of the cowling and cooling of the ranger V-770-8 engine installation in the Edo XOSE-1 airplane. Three types of baffles were installed in the course of the tests: the conventional, the turbulent-flow, and the NACA diffuser baffles. Each of the types was of merit in cooling a different region on the cylinder. Incorporation of the best features of the three types into one baffle, a method which appears to be feasible, would provide improvements in cylinder cooling.

COOLING TESTS↗
NASA NTRS1954

A Method of Measuring Jet Thrust of Turbojet Engines in Flight Installations

Measurement of the jet thrust of a turbojet engine in flight becomes more difficult as the number and complexity of the engine components increase. It is desirable, therefore, that a general correlation of jet thrust be developed which is applicable to a simple direct reading thrustmeter. In view of this need a correlation is presented which is independent of flight conditions and applicable to both non-afterburning and afterburning engines equipped with nonejector type fixed- and variable-area convergent exhaust nozzles. The general equation used in this correlation was derived from the theoretical jet-thrust equation for a choked convergent nozzle. The data used to verify the correlation were obtained over a range of altitudes from 10,000 to 54,000 feet and a range of flight Mach numbers from 0.4 to 1.1. A thrustometer based on this equation was installed on an afterburning turbojet engine equipped with a fixed area convergent exhaust nozzle. The results indicated that a meter based on this correlation would be applicable to a flight installation and that the probable error in thrust measurement would be approximately + or -1.5 percent provided the exhaust-nozzle thrust coefficient is known.

Sivo, Joseph N↗
NASA NTRS1994

Blind-Anchor-Nut-Installation Fixture (BANIF)

Blind-anchor-nut-installation fixture, BANIF, developed for replacing or installing anchor nuts in blind holes or other inaccessible places. Attachment of anchor nut to BANIF enables placement of anchor nut on blind side of component.

Willey, Norman F., Jr.↗
NASA NTRS1994

MSFC inspections of installed polyimide wire

An alert was issued because of the arc-tracking possibilities of installed polyimide wire harnesses. MSFC undertook a program to try to enhance the safety and reliability of these harnesses. Photographs are presented showing the need for inspections of installed wiring harnesses.

Landers, Joe C.↗
NASA NTRS1994

Installed F/A-18 inlet flow calculations at 30 degrees angle-of-attack: A comparative study

NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high-performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Full Navier-Stokes (FNS) calculations on the installed F/A-18 inlet at 30 degrees angle of attack, 0 degrees yaw, and a freestream Mach number of 0.2 have been obtained in this study using an algebraic turbulence model with two grids (original and revised). Results obtained with the original grid were used to determine where further grid refinements and additional geometry were needed. In order to account properly for the external effects, the forebody, leading edge extension (LEX), ramp, and wing were included with inlet geometry. In the original grid, the diverter, LEX slot, and leading edge flap were not included due to insufficient geometry definition, but were included in a revised grid. In addition, a thin-layer Navier-Stokes (TLNS) code is used with the revised grid and the numerical results are compared to those obtained with the FNS code. The TLNS code was used to evaluate the effects on the solution using a code with more recent CFD developments such as upwinding with TVD schemes versus central differencing with artificial dissipation. The calculations are compared to a limited amount of available experimental data. The predicted forebody/fuselage surface static pressures compared well with data of all solutions. The predicted trajectory of the vortex generated under the LEX was different for each solution. These discrepancies are attributed to differences in the grid resolution and turbulence modeling. All solutions predict that this vortex is ingested by the inlet. The predicted inlet total pressure recoveries are lower than data and the distortions are higher than data. The results obtained with the revised grid were significantly improved from the original grid results. The original grid results indicated the ingested vortex migrated to the engine face and caused additional distortions to those already present due to secondary flow development. The revised grid results indicate that the ingested vortex is dissipated along the inlet duct inboard wall. The TLNS results indicate the flow at the engine face was much more distorted than the FNS results and is attributed to the pole boundary condition introducing numerical distortions into the flow field.

Smith, C. Frederic↗
NASA NTRS1959

Review of Aircraft Altitude Errors Due to Static-Pressure Source and Description of Nose-Boom Installations for Aerodynamic Compensation of Error

A brief review of airplane altitude errors due to typical pressure installations at the fuselage nose, the wing tip, and the vertical fins is presented. A static-pressure tube designed to compensate for the position errors of fuselage-nose installations in the subsonic speed range is described. This type of tube has an ogival nose shape with the static-pressure orifices located in the low-pressure region near the tip. The results of wind-tunnel tests of these compensated tubes at two distances ahead of a model of an aircraft showed the position errors to be compensated to within 1/2 percent of the static pressure through a Mach number range up to about 1.0. This accuracy of sensing free-stream static pressure was extended up to a Mach number of about 1.15 by use of an orifice arrangement for producing approximate free-stream pressures at supersonic speeds and induced pressures for compensation of error at subsonic speeds.

Gracey, William↗
NASA NTRS1960

Measurement of the Errors of Service Altimeter Installations During Landing-Approach and Take-Off Operations

The overall errors of the service altimeter installations of a variety of civil transport, military, and general-aviation airplanes have been experimentally determined during normal landing-approach and take-off operations. The average height above the runway at which the data were obtained was about 280 feet for the landings and about 440 feet for the take-offs. An analysis of the data obtained from 196 airplanes during 415 landing approaches and from 70 airplanes during 152 take-offs showed that: 1. The overall error of the altimeter installations in the landing- approach condition had a probable value (50 percent probability) of +/- 36 feet and a maximum probable value (99.7 percent probability) of +/- 159 feet with a bias of +10 feet. 2. The overall error in the take-off condition had a probable value of +/- 47 feet and a maximum probable value of +/- 207 feet with a bias of -33 feet. 3. The overall errors of the military airplanes were generally larger than those of the civil transports in both the landing-approach and take-off conditions. In the landing-approach condition the probable error and the maximum probable error of the military airplanes were +/- 43 and +/- 189 feet, respectively, with a bias of +15 feet, whereas those for the civil transports were +/- 22 and +/- 96 feet, respectively, with a bias of +1 foot. 4. The bias values of the error distributions (+10 feet for the landings and -33 feet for the take-offs) appear to represent a measure of the hysteresis characteristics (after effect and recovery) and friction of the instrument and the pressure lag of the tubing-instrument system.

Gracey, William↗
NASA NTRS1999

Delta FUSE Fairing Installation at Launch Complex 17A

This NASA Kennedy Space Center (KSC) video release presents footage of the June 19, 1999 installation of the fairing around the Far Ultraviolet Spectroscopic Explorer (FUSE) spacecraft. The spacecraft was previously mated to the Boeing Delta II rocket. Installation took place on Pad A of Launch Complex 17.

Source record↗
NASA NTRS1999

Summary of HEAT 1 Aeroacoustics Installation Effects

A critical part of the NASA High-Speed Research (HSR) program is the demonstration of satisfactory suppression of the jet noise present at low airspeeds. One scheme for reducing jet exhaust noise generated by a future High-Speed Civil Transport (HSCT) is the use of a mixer/ ejector system which would entrain large quantities of ambient air into the exhaust flow from the powerplant in order to cool and slow the jet exhaust before it leaves the tailpipe. Of the variety of factors which can affect the noise suppression characteristics of the mixer/ejector system, the influence of the wing flow field and high-lift devices is not well understood. The effectiveness of the noise suppression device must be evaluated in the presence of the wing/high-lift system before definitive assessments can be made concerning HSCT noise. Of nearly equal importance is the evaluation of the performance of the high-lift system(s) in the presence of realistic propulsion units which feature high ambient flow entrainment rates and jet thrust coefficients. These noise suppressors must provide the required acoustic attenuation while not overly degrading the thrust efficiency of the propulsion system or the lift enhancement of the high-lift devices on the wing. The overall objective of the NASA High-lift Engine Aeroacoustics Technology program is to demonstrate satisfactory interaction between the jet noise suppressor and the high-lift system at airspeeds and angles of attack consistent with takeoff, climb, approach, and landing. In support of this program, an isolated aeroacoustic test of a 13.5%-scale, candidate mixer/ejector nozzle was performed in the Ames' Research Center 40- by 80-Foot Wind Tunnel. The purpose of the test was to measure the baseline aeroacoustic performance characteristics of this nozzle in isolation from the aerodynamic flowfield induced by an HSCT airframe. The test documented the acoustic signature of the nozzles with treated and hardwall ejector surfaces and with changes in the ratio of ejector-duct-to-jet-area over a wide range range of nozzle pressure ratios and freestream Mach numbers. The test also measured the thrust performance, ambient-flow aspiration ratio, and internal and external static pressures on the nozzles. The isolated aeroacoustic performance data has been compared with results obtained with this nozzle installed on a 13.5% Boeing Reference H HSCT configuration, semi-span model. The semi-span, aeroacoustics integration test documented the first-order effects of the airframe flowfield on the acoustic performance of the nozzle and the effect of the nozzle secondary inlet flows on the aerodynamic performance of the wing high-lift systems. This investigation is critical to understanding the mutual installation effects of mixer/ejector nozzles and wing high-lift systems.

Smith, Brian E.↗
NASA NTRS2001

Computational Analyses of Propulsion Aeroacoustics for Mixed Flow Nozzle Pylon Installation at Takeoff

A CFD analyses is presented for a set of baseline and noise suppression mixed flow nozzles with and without a pylon installation. The five model configurations are as follows; a baselinecore/fan dual-stream nozzle with an external plug, a chevron mixer nozzle with a peak on the symmetry plane with external plug, both of the above nozzles with an installed bifurcatingpylon and lastly a clocked chevron mixer nozzle such that a trough is aligned with the center of the pylon. The fluid flow is simulated by solving the asymptotically steady, compressible, Reynolds-averaged Navier-Stokes equations using an implicit, up-wind, flux-difference splitting finite volume scheme and standard two equation k-epsilon turbulence model with a linear stress representation. All computations are performed using the multiblock, parallel, structuredcode PAB3D. Results indicate that the clocked chevron with pylon case achieves the most optimal levels of average and peak turbulence kinetic energy and vorticity and therefore is expected to be the quietest of the five configurations tested. Further study is required to refine expressions which are indicative of noise and mate these with rigorous noise prediction models.

Massey, Steven J.↗
NASA NTRS2004

Suggested Procedures for Installing Strain Gauges on Langley Research Center Wind Tunnel Balances, Custom Force Measuring Transducers, Metallic and Composite Structural Test Articles

The character of force and strain measurement testing at LaRC is such that the types of strain gauge installations, the materials upon which the strain gauges are applied, and the test environments encountered, require many varied approaches. In 1997, a NASA Technical Memorandum (NASA TM 110327) was generated to provide the strain gauge application specialist with a listing of recommended procedures for strain gauging various transducers and test articles at LaRC. The technical memorandum offered here is an effort to keep the strain gauge user informed of new technological enhancements in strain-gauging methodology while preserving the strain-gauging guidelines set forth in the 1997 TM. This document provides detailed recommendations for strain gauging LaRC-designed balances and custom transducers, composite materials, cryogenic and high-temperature test articles, and selected non-typical or unique materials or test conditions. Additionally, one section offers details for installing Bragg-Grating type fiber-optic strain sensors for non-typical test scenarios.

Moore, Thomas C., Sr.↗
NASA NTRS1998

Effect of Installation of Mixer/Ejector Nozzles on the Core Flow Exhaust of High-Bypass-Ratio Turbofan Engines

The aerospace industry is currently investigating the effect of installing mixer/ejector nozzles on the core flow exhaust of high-bypass-ratio turbofan engines. This effort includes both full-scale engine tests at sea level conditions and subscale tests in static test facilities. Subscale model tests are to be conducted prior to full-scale testing. With this approach, model results can be analyzed and compared with analytical predications. Problem areas can then be identified and design changes made and verified in subscale prior to committing to any final design configurations for engine ground tests. One of the subscale model test programs for the integrated mixer/ejector development was a joint test conducted by the NASA Lewis Research Center and Pratt & Whitney Aircraft. This test was conducted to study various mixer/ejector nozzle configurations installed on the core flow exhaust of advanced, high-bypass-ratio turbofan engines for subsonic, commercial applications. The mixer/ejector concept involves the introduction of largescale, low-loss, streamwise vortices that entrain large amounts of secondary air and rapidly mix it with the primary stream. This results in increased ejector pumping relative to conventional ejectors and in more complete mixing within the ejector shroud. The latter improves thrust performance through the efficient energy exchange between the primary and secondary streams. This experimental program was completed in April 1997 in Lewis' CE-22 static test facility. Variables tested included the nozzle area ratio (A9/A8), which ranged from 1.6 to 3.0. This ratio was varied by increasing or decreasing the nozzle throat area, A8. Primary nozzles tested included both lobed mixers and conical primaries. These configurations were tested with and without an outer shroud, and the shroud position was varied by inserting spacers in it. In addition, data were acquired with and without secondary flow.

Harrington, Douglas E.↗
NASA NTRS2003

Static Frequency Converter System Installed and Tested

A new Static Frequency Converter (SFC) system has been installed and tested at the NASA Glenn Research Center s Central Air Equipment Building to provide consistent, reduced motor start times and improved reliability for the building s 14 large exhausters and compressors. The operational start times have been consistent around 2 min, 20 s per machine. This is at least a 3-min improvement (per machine) over the old variable-frequency motor generator sets. The SFC was designed and built by Asea Brown Boveri (ABB) and installed by Encompass Design Group (EDG) as part of a Construction of Facilities project managed by Glenn (Robert Scheidegger, project manager). The authors designed the Central Process Distributed Control Systems interface and control between the programmable logic controller, solid-state exciter, and switchgear, which was constructed by Gilcrest Electric.

Brown, Donald P.↗
NASA NTRS2004

Seal ring installation tool

A seal ring tool that allows an installer to position a primary seal ring between hub ends of pipe flanges that are being assembled together. The tool includes a pivoting handle member and extension arms attached to the pivoting handle member. The ends of the arms have side indentation type longitudinal grooves angled toward one another for holding the primary seal ring in place between the hubs of respective pipes that are to be attached together. The arms of the tool can also have flat sides that can be used to abut against an optional second larger seal that is supported within a groove in one of the hub ends so that the second hub end can then be moved against the other side of the primary seal ring. Once the seal ring is positioned between the pipe hubs, the pipe hubs can be moved about the seal ring due to the flat sides of the arms of the tool. The tool eliminates the chances of damaging and contaminating seal rings being installed within pipe hubs that are being attached to one another.

Haselmaier, L. Haynes↗
NASA NTRS1939

Radiator Design and Installation

The fundamental principles of fluid flow, pressure losses, and heat transfer have been presented and analyzed for the case of a smooth tube with fully developed turbulent flow. These equations apply to tubes with large length-diameter ratios where the f1ow is at a high Reynolds Number. The error introduced by using these equations increases as the magnitude of the tube length and the air-flow Reynolds Number approaches the values encountered in modern radiator designs. Accordingly, heat-transfer tests on radiator sections were made and the results are presented in nondimensional form to facilitate their use and for comparison with other heat-transfer data. In addition, pressure losses were measured along smooth tubes of circular, square, and rectangular cross section and the results were also correlated and are presented in nondimensional form. The problem of a radiator design for a particular installation is solved, the experimental heat-transfer and pressure-loss data being used, on a basis of power chargeable to the radiator for form drag, for propelling the weight, and for forcing the air through the radiator. The case of an installation within a wing or an engine nacelle is considered. An illustration of radiator design is carried through for an arbitrary set of conditions. Sufficient detail is given to enable the reader to reproduce the analysis for any given case.

Brevoort, M.J.↗