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Full scale wind tunnel tests of low wing, single engine, light aircraft with positive and negative propeller thrust and up and down flap deflection - graphs
A sharp leading edge delta-wing model was tested through an angle-of-attack range at Mach numbers of 0.75, 0.80, and 0.85 at both ambient and cryogenic temperatures in the Langley 1/3-meter transonic cryogenic tunnel. Total pressure was varied with total temperature in order to hold test Reynolds number constant at a given Mach number. Agreement between the aerodynamic data obtained at ambient and cryogenic temperatures indicates that flows with leading-edge vortex effects are duplicated properly at cryogenic temperatures. The test results demonstrate that accurate aerodynamic data can be obtained by using conventional force-testing techniques if suitable measures are taken to minimize temperature gradients across the balance and to keep the balance at ambient (warm) temperatures during cryogenic operation of the tunnel.
Four six-place, low-wing, twin-engine, general aviation airplane test specimens were crash tested under controlled free flight conditions. All airplanes were impacted on a concrete test surface at a nomial flight path velocity of 27 m/sec. Two tests were conducted at a -15 deg flight path angle (0 deg pitch angle and 15 deg pitch angle), and two were conducted at a -30 deg flight path angle (-30 deg pitch angle). The average acceleration time histories (crash pulses) in the cabin area for each principal direction were calculated for each crash test. In addition, the peak floor accelerations were calculated for each test as a function of aircraft fuselage longitudinal station number. Anthropomorphic dummy accelerations were analyzed using the dynamic response index and severity index (SI) models. Parameters affecting the dummy restraint system were studied; these parameters included the effect of no upper torso restraint, measurement of the amount of inertia-reel strap pullout before locking, measurement of dummy chest forward motion, and loads in the restraints. With the SI model, the dummies with no shoulder harness received head impacts above the concussive threshold.
This paper investigates the utility of the Hilbert-Huang transform for the analysis of aeroelastic flight data. It is well known that the classical Hilbert transform can be used for time-frequency analysis of functions or signals. Unfortunately, the Hilbert transform can only be effectively applied to an extremely small class of signals, namely those that are characterized by a single frequency component at any instant in time. The recently-developed Hilbert-Huang algorithm addresses the limitations of the classical Hilbert transform through a process known as empirical mode decomposition. Using this approach, the data is filtered into a series of intrinsic mode functions, each of which admits a well-behaved Hilbert transform. In this manner, the Hilbert-Huang algorithm affords time-frequency analysis of a large class of signals. This powerful tool has been applied in the analysis of scientific data, structural system identification, mechanical system fault detection, and even image processing. The purpose of this paper is to demonstrate the potential applications of the Hilbert-Huang algorithm for the analysis of aeroelastic systems, with improvements such as localized/online processing. Applications for correlations between system input and output, and amongst output sensors, are discussed to characterize the time-varying amplitude and frequency correlations present in the various components of multiple data channels. Online stability analyses and modal identification are also presented. Examples are given using aeroelastic test data from the F/A-18 Active Aeroelastic Wing aircraft, an Aerostructures Test Wing, and pitch-plunge simulation.
The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.
This report describes strain-gage calibration loading through the application of known loads of the Active Aeroelastic Wing F/A-18 airplane. The primary goal of this test is to produce a database suitable for deriving load equations for left and right wing root and fold shear; bending moment; torque; and all eight wing control-surface hinge moments. A secondary goal is to produce a database of wing deflections measured by string potentiometers and the onboard flight deflection measurement system. Another goal is to produce strain-gage data through both the laboratory data acquisition system and the onboard aircraft data system as a check of the aircraft system. Thirty-two hydraulic jacks have applied loads through whiffletrees to 104 tension-compression load pads bonded to the lower wing surfaces. The load pads covered approximately 60 percent of the lower wing surface. A series of 72 load cases has been performed, including single-point, double-point, and distributed load cases. Applied loads have reached 70 percent of the flight limit load. Maximum wingtip deflection has reached nearly 16 in.
Wind tunnel tests of blunt reentry vehicle with retractable wings
This report presents the results of an exploratory investigation carried out in the Langley 9-inch supersonic tunnel to determine the maximum lift of wings operating at supersonic speeds. A variety of wing plan forms of random thickness distributions were tested at Mach numbers of 1.55, 1.90, and 2.32 and at Reynolds numbers varying between 0.74 x 10(6) and 0.27 x 10(6) at angles of attack ranging from zero up through the angle at which maximum lift occurred. Subsequent pressure-distribution tests on wings of triangular and rectangular plan forms were made at a Mach number of 2.40. The results of these tests substantiated the values of maximum lift obtained during the force tests and further showed no appreciable center-of-pressure shift over the entire angle-of-attack range.
The Ames 12-Foot Pressure Tunnel was used to determine the effects of Reynolds number on the static longitudinal aerodynamic characteristics of an advanced, high-aspect-ratio, supercritical wing transport model equipped with a full span, leading edge slat and part span, double slotted, trailing edge flaps. The model had a wing span of 7.5 ft and was tested through a free stream Reynolds number range from 1.3 to 6.0 x 10 to 6th power per foot at a Mach number of 0.20. Prior to the Ames tests, an investigation was also conducted in the Langley 4 by 7 Meter Tunnel at a Reynolds number of 1.3 x 10 to 6th power per foot with the model mounted on an Ames strut support system and on the Langley sting support system to determine strut interference corrections. The data obtained from the Langley tests were also used to compare the aerodynamic charactertistics of the rather stiff, 7.5-ft-span steel wing model tested during this investigation and the larger, and rather flexible, 12-ft-span aluminum-wing model tested during a previous investigation. During the tests in both the Langley and Ames tunnels, the model was tested with six basic wing configurations: (1) cruise; (2) climb (slats only extended); (3) 15 deg take-off flaps; (4) 30 deg take-off flaps; (5) 45 deg landing flaps; and (6) 60 deg landing flaps.
The paper continues the status reporting of the ultralightweight flexible blanket, flatpack, foldout solar array testbed wing that was presented at the previous Meeting. The test bed wing has been built and subjected to a variety of critical functional tests after exposure to simulated launch environments.
Free flight tests of 45-deg swept wings of 3.15 aspect ratio and 0.54 taper ratio to measure wing damping and possible transonic flutter
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Noise measurements were made to determine if recent modifications made to the bifurcated jetpipe to increase engine thrust had at the same time reduced the noise level. The noise field was measured by a 6-microphone array positioned on a 30.5m (100 ft) sideline between 90 and 150 degrees from the left engine inlet. Noise levels were recorded at three flap angles over a range of engine thrust settings from flight idle to emergency power and plotted in one-third octave band spectra. Little attenuation was observed at maximum power, but significant attenuation was achieved at approach and cruise power levels.
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Explore the source record for details and available documents.
Explore the source record for details and available documents.
Explore the source record for details and available documents.