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NASA NTRS2011

Pressure-Sensitive Paint Measurements on the NASA Common Research Model in the NASA 11-ft Transonic Wind Tunnel

The luminescence lifetime technique was used to make pressure-sensitive paint (PSP) measurements on a 2.7% Common Research Model in the NASA Ames 11ft Transonic Wind Tunnel. PSP data were obtained on the upper and lower surfaces of the wing and horizontal tail, as well as one side of the fuselage. Data were taken for several model attitudes of interest at Mach numbers between 0.70 and 0.87. Image data were mapped onto a three-dimensional surface grid suitable both for comparison with CFD and for integration of pressures to determine loads. Luminescence lifetime measurements were made using strobed LED (light-emitting diode) lamps to illuminate the PSP and fast-framing interline transfer cameras to acquire the PSP emission.

Bell, James H.↗

NASA NTRSDate not supplied

A Summary of Test Results from a NASA Lift + Cruise eVTOL Crash Test

On November 9, 2022, the National Aeronautics and Space Administration (NASA) conducted a full-scale crash test of the NASA Lift+Cruise (LPC) reference vehicle at the NASA Langley Research Center Landing and Impact Research Facility (LandIR) under combined vertical and horizontal impact conditions to simulate a severe but survivable crash. The LPC test article is a carbon-composite skin/frame structure design, developed and fabricated for the cabin section only. The test utilized various configurations of seats and Anthropomorphic Test Devices (ATDs, a.k.a. crash test dummies) intended to encompass a variety of occupant conditions. In addition, an in-house developed energy absorbing subfloor was utilized for the evaluation of load attenuation. Overhead mass was simulated using attached lifting hardware and other systems were simulated using ballast mass. The test article impacted the ground with velocities of 38.1 ft/s horizontal and 31.4 ft/s vertical. During the first approximately 38 milliseconds (ms), the cabin section experienced a large amount of acceleration on the belly which was attenuated by the subfloor structures and seats. Over the next approximately 160 ms, the test article experienced failure in the a-, b- and c-pillars, leading to a partial collapse of the overhead structure. Test data was collected on the belly, floors, seats, overhead mass, and tail. In addition, there was high speed full-field digital image correlation analysis data collected on the port side skin surface. Collected ATD data showed that the ATDs seated in the energy absorbing seats experienced loads at limits below the those in current regulations. Even without an energy absorbing seat, the energy absorbing subfloors crushed as intended, and limited the load on the large 95 th percentile ATD to 10% over suggested limits. The collapse of the roof did affect the 95 th percentile ATD, which showed high neck loading values due to head contact, whereas survivable volume was maintained for the other ATDs. While the energy absorbing subfloor and energy absorbing seats both contributed to occupant load attenuation, further optimization is suggested to increase their robustness.

evtol↗

NASA NTRSDate not supplied

A Summary of Test Results from a NASA Lift + Cruise eVTOL Crash Test

evtol↗

NASA NTRS1974

Subsonic and supersonic longitudinal stability and control characteristics of an aft tail fighter configuration with cambered and uncambered wings and uncambered fuselage

An investigation has been made in the Mach number range from 0.20 to 2.16 to determine the longitudinal aerodynamic characteristics of a fighter airplane concept. The configuration concept employs a single fixed geometry inlet, a 50 deg leading-edge-angle clipped-arrow wing, a single large vertical tail, and low horizontal tails. The wing camber surface was optimized in drag due to lift and was designed to be self-trimming at Mach 1.40 and at a lift coefficient of 0.20. An uncambered or flat wing of the same planform and thickness ratio was also tested. However, for the present investigation, the fuselage was not cambered. Further tests should be made on a cambered fuselage version, which attempts to preserve the optimum wing loading on that part of the theoretical wing enclosed by the fuselage.

Dollyhigh, S. M.↗

NASA NTRS1990

Feasibility study for a microwave-powered ozone sniffer aircraft

The preliminary design of a high-altitude, remotely-piloted, atmospheric-sampling aircraft powered by microwave energy beamed from ground-based antenna was completed. The vehicle has a gross weight of 6720 pounds and is sized to carry a 1000 pound payload at an altitude of 100,000 feet. The underside of the wing serves as the surface of a rectenna designed to receive microwave energy at a power density of 700 watts per square meter and the wing has a planform area of 3634 square feet to absorb the required power at an optimum Mach number M = 0.44. The aircraft utilizes a horizontal tail and a canard for longitudinal control and to enhance the structural rigidity of the twin fuselage configuration. The wing structure is designed to withstand a gust-induced load factor n = 3 at cruise altitude but the low-wing loading of the aircraft makes it very sensitive to gusts at low altitudes, which may induce load factors in excess of 20. A structural load alleviation system is therefore proposed to limit actual loads to the designed structural limit. Losses will require transmitted power on the order of megawatts to be radiated to the aircraft from the ground station, presenting environmental problems. Since the transmitting antenna would have a diameter of several hundred feet, it would not be readily transportable, so we propose that a single antenna be constructed at a site from which the aircraft is flown. The aircraft would be towed aloft to an initial altitude at which the microwave power would be utilized. The aircraft would climb to cruise altitude in a spiral flight path and orbit the transmitter in a gentle turn.

Botros, David F.↗

NASA NTRS2007

Plans and Status of Wind-Tunnel Testing Employing an Aeroservoelastic Semispan Model

This paper presents the research objectives, summarizes the pre-wind-tunnel-test experimental results to date, summarizes the analytical predictions to date, and outlines the wind-tunnel-test plans for an aeroservoelastic semispan wind-tunnel model. The model is referred to as the Supersonic Semispan Transport (S4T) Active Controls Testbed (ACT) and is based on a supersonic cruise configuration. The model has three hydraulically-actuated surfaces (all-movable horizontal tail, all-movable ride control vane, and aileron) for active controls. The model is instrumented with accelerometers, unsteady pressure transducers, and strain gages and will be mounted on a 5-component sidewall balance. The model will be tested twice in the Langley Transonic Dynamics Tunnel (TDT). The first entry will be an "open-loop" model-characterization test; the second entry will be a "closed-loop" test during which active flutter suppression, gust load alleviation and ride quality control experiments will be conducted.

Perry, Boyd, III↗

NASA NTRS1940

Pressure-distribution investigation of an NACA 0009 airfoil with an 80-percent-chord plain flap and three tabs

Pressure-distribution tests of an NACA 0009 airfoil with an 80-percent-chord plain flap and three plain tabs, having chord of 10, 20, and 30 percent of the flap chord, were made. Section data suitable for application to the design of horizontal and vertical tail surfaces were obtained. Resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab are presented. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. A comparison of some characteristic slopes for the 30-, the 50-, and the 80-percent-chord flaps, tested in the general investigation of plain flaps for control surfaces, is included. Section aerodynamic and load data have been made available for a wide range of flap and a tab chords to be used on an NACA 0009 airfoil or on other conventional sections.

Ames, Milton B , Jr↗

NASA NTRS1993

The airplane: A simulated commercial air transportation study

The 'Airplane' is a moderate-range, 70 passenger aircraft. It is designed to serve demands for flights up to 10,000 feet and it cruises at 32 ft/s. The major drivers for the design of the Airplane are economic competitiveness, takeoff performance, and weight minimization. The Airplane is propelled by a single Astro 15 electric motor and a Zinger 12-8 propeller. The wing section is a Spica airfoil which, because of its flat bottom, provides simplicity in manufacturing and thus helps to cut costs. The wing is constructed of a single load bearing mainspar and shape-holding ribs coated with Monokote skin, lending to a light weight structural makeup. The fuselage houses the motor, flight deck and passenger compartments as well as the fuel and control actuating systems. The wing will be attached to the top of the fuselage as will the fuel and control actuator systems for easy disassembly and maintenance. The aircraft is maneuvered about its pitch axis by means of an aft elevator on the flat plate horizontal tail. The twin vertical tail surfaces are also flat plates and each features a rudder for both directional and roll control. Along with wing dihedral, the rudders will be used to roll the aircraft. The Airplane is less costly to operate at its own maximum range and capacity as well as at its maximum range and the HB-40's maximum capacity than the HB-40.

Dauteuil, Mark↗

NASA NTRS1946

Free-spinning-tunnel Tests of a 1/26th Scale Model of the Douglas XTB2D-1 Airplane

A spin-tunnel investigation of a 1/26 scale model of the Douglas XTB2D-1 airplane has been conducted in the Langley 20-foot free-spinning tunnel. The effects of control settings and movements upon the erect- and inverted-spin and recovery characteristics of the model were determined for various loading conditions. Tests were also performed to determine the effects of various tail modifications. The investigation included emergency spin-recovery parachute tests as well as crew-escape and rudder- and elevator-force tests. All tests were performed at an equivalent spin altitute of 20,000 feet. The recovery characteristics of the model in the original design were found to be unsatisfactory. Installation of a large ventral fin, installation of tip fins on the horizontal tail, or installation of a small ventral fin in combination with antispan fillets and a spanwise extension of the horizontal-tail surfaces satisfactorily improved the recovery characteristics of the model.

Stone, Ralph W↗

NASA NTRS1937

Pressure-Distribution Measurements on O-2H Airplane in Flight

Results are given of pressure-distribution measurements made over two different horizontal tail surfaces and the right wing cellule, including the slipstream area, of an observation-type biplane. Measurements were also taken of air speed, control-surface positions, control-stick forces, angular velocities, and accelerations during various abrupt maneuvers. These maneuvers consisted of push-downs and pull-ups from level flight, dive pull-outs, and aileron rolls with various thrust conditions. The results from the pressure-distribution measurements over the wing cellule are given on charts showing the variation of individual rib coefficients with wing coefficients; the data from the tail-surface pressure-distribution measurements are given mainly as total loads and moments. These data are supplemented by time histories of the measured quantities and isometric views of the rib pressure distributions occurring in abrupt maneuvers.

Pearson, H A↗

NASA NTRS1938

Flight and Wind-tunnel Tests of an XBM-1 Dive Bomber

Results are given of pressure-distribution measurements made in flight over the right wing cellule and the right half of the horizontal tail surfaces of a dive-bombing biplane. Simultaneous measurements were also taken of the air speed, control-surface positions, control forces, and normal accelerations during various abrupt maneuvers in vertical plane. These maneuvers consisted of push-downs and pull-ups from level flight, dives and dive pull-ups from inverted flight. Besides the pressure measurements, flight tests were made to obtain (1) wing-fabric deflections during dives and (2) variation of the minimum drag coefficient with Reynolds Number. Supplementary tests were also done in the full-scale wind tunnel to obtain the characteristics of the airplane under various propeller conditions and with various tail settings. The results indicate that: (1) by increasing the fabric deflection between pressure ribs, the span load distribution was considerably modified near the center and the wing moment relations were changed; and (2) the minimum drag was less for the idling propeller than for the propeller locked in a vertical position. The value of C(sub D sub min) was equal to K(Reynolds Number)(exp -0.03) for a range from 2,800,000 to 13,100,000.

Donely, Philip↗

NASA NTRS1954

Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing-tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing-tail interference over the angle-of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg. The wing-tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the crossflow drag coefficient along the body and the effect of the wing downwash field on the afterbody loading.

Spahr, J. R.↗

NASA NTRS1954

Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing- tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift - interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing- tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing- tail interference over the angle - of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg center dot The wing- tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the cross flow drag coefficient along the body and the effect of the wing downwash field on the after body loading.

Spahr, J Richard↗